diff --git "a/logs/app.log" "b/logs/app.log" --- "a/logs/app.log" +++ "b/logs/app.log" @@ -25301,3 +25301,50159 @@ Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report A ------ 2025-04-04 at 04:03:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:03:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "L1 azimuth" or "vehicle azimuth at liftoff of a Falcon 9 +2025-04-04 at 04:03:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:03:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 17 launch azimuth at t=0 Apollo 12 navigation accuracy +2025-04-04 at 04:03:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:03:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch azimuth +2025-04-04 at 04:03:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:03:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:03:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: azimuth at liftoff of a succesful Falcon 9 launch +2025-04-04 at 04:03:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:03:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch azimuth November 1970 +2025-04-04 at 04:03:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:03:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:03:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Falcon 9 L1 launch azimuth +2025-04-04 at 04:03:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:03:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch vehicle azimuth angle +2025-04-04 at 04:03:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-04 at 04:03:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:03:48 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:03:48 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:03:48 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, True, True, True, False] +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.67 +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_correctness:82 - Student lengths: [482, 803, 338, 199, 26, 261] +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [24, 24, 24, 24, 24, 24] +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_correctness:84 - Average student length: 351.50 +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 24.00 +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_correctness:86 - Length ratio: 14.65 +2025-04-04 at 04:03:48 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-04 at 04:03:48 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.508 Âą 0.412 +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.83 Âą 2.61 +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 1, 6, 4, 6, 0] +2025-04-04 at 04:03:48 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:03:48 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...'] +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-04 at 04:03:48 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...'] +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-04 at 04:03:48 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 3.0 +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.500 +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.500 +2025-04-04 at 04:03:48 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.267, Perfect scores: 1/6 +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 2/6 +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.33 +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:03:48 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:03:48 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.320, Max reward: 0.583 +2025-04-04 at 04:03:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:03:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the first balloons to reach the edge of space? +2025-04-04 at 04:03:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:03:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Montgolfier brothers' 1783 hot air balloon launch time +2025-04-04 at 04:03:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:03:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"when did the first helium filled balloon launch occur" +and +"when did the first high altitude balloon launch occur +2025-04-04 at 04:03:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ + +2025-04-04 at 04:03:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "First two balloons to go to space +2025-04-04 at 04:03:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:03:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:03:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first balloons orbit geostationary altitude +2025-04-04 at 04:03:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:03:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 launch time +2025-04-04 at 04:03:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:03:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: type "NASA first two balloons launched in space +2025-04-04 at 04:03:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:03:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:03:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: James A. Lovell, Fred W. Haise, and John L. Swigert Apollo 11 launch time +2025-04-04 at 04:03:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-04 at 04:03:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 schedule launch +2025-04-04 at 04:03:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:03:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:04:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch time +2025-04-04 at 04:04:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:04:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:04:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch time April 12 +2025-04-04 at 04:04:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:04:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:04:03 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:04:03 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:04:03 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, True, False] +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_correctness:82 - Student lengths: [184, 513, 426, 703, 327, 756] +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_correctness:84 - Average student length: 484.83 +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 13.00 +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_correctness:86 - Length ratio: 37.29 +2025-04-04 at 04:04:03 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-04 at 04:04:03 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.158 Âą 0.354 +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 0.83 Âą 1.86 +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 0, 0, 5, 0] +2025-04-04 at 04:04:03 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ + +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-04 at 04:04:03 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-04 at 04:04:03 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:04:03 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 0/6 +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 0.00 +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 0/6 +2025-04-04 at 04:04:03 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:04:03 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.326, Max reward: 0.627 +2025-04-04 at 04:04:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:04:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "y-axis error magnitude discrepancy analysis" or " experimental data error source analysis +2025-04-04 at 04:04:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-04 at 04:04:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the Y-axis error discrepancy in nuclear thermal-hydraulic analysis +2025-04-04 at 04:04:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-04 at 04:04:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:04:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "gyrocompassing misalignment null bias drift" or "ifax inertial component preflight histories +2025-04-04 at 04:04:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-04 at 04:04:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: why did a spacecraft encounter a Y-axis error +2025-04-04 at 04:04:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ +Result 2: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ + +2025-04-04 at 04:04:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:04:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module platform gyro alignment difficulties" or "gyroscopic alignment techniques in space +2025-04-04 at 04:04:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-04 at 04:04:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: attitude error calculation initialization in reentry maneuver +2025-04-04 at 04:04:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-04 at 04:04:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:04:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module to command module alignment problem" +2025-04-04 at 04:04:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-04 at 04:04:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: actually forged attitudes with in ship fly-by-wire system +2025-04-04 at 04:04:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ + +2025-04-04 at 04:04:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:04:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module undocking maneuver difficulties" or "command module gimbal lock problem +2025-04-04 at 04:04:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-04 at 04:04:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what were the implications of the out-of-sequence turn-on procedure on the spacecraft's navigation +2025-04-04 at 04:04:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-04 at 04:04:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:04:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module undocking optical alignment difficulties" or "coarse alignment to ground-supplied reference angles +2025-04-04 at 04:04:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:04:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:04:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "manually maneuvering to escape trajectory alignment" or "contingency planning for star alignment +2025-04-04 at 04:04:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:04:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:04:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "manually aligning spacecraft with body axes" or " Abort Guidance System operation +2025-04-04 at 04:04:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:04:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:04:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "inertial measurement accuracy corrections" or "star-angle data errors +2025-04-04 at 04:04:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:04:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:04:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Abort Guidance System performance summary" or "inertial accuracy analysis +2025-04-04 at 04:04:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:04:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:04:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "transearth injection attitude analysis" or "coast phase alignment summary +2025-04-04 at 04:04:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:04:25 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:04:25 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:04:25 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1605, 588, 495, 663, 280, 1485] +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [36, 36, 36, 36, 36, 36] +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_correctness:84 - Average student length: 852.67 +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 36.00 +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_correctness:86 - Length ratio: 23.69 +2025-04-04 at 04:04:25 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-04 at 04:04:25 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.000 Âą 0.000 +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 0.00 Âą 0.00 +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 0, 0, 0, 0] +2025-04-04 at 04:04:25 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-04 at 04:04:25 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-04 at 04:04:25 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-04 at 04:04:25 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-04 at 04:04:25 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ +Result 2: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ + +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ + +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-04 at 04:04:25 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nError Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Fliph...', 'Result 1:\nThe error counters for the coupling display units are used by the digital aut...', 'Result 1:\nAfter completing the maneuver to the attitude required for the final midcours...', 'Result 1:\nstarted to sight the service module in the docking window. The lightened spac...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...'] +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 1.0 +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.167 +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:04:25 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 1/6 +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 0.50 +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 0/6 +2025-04-04 at 04:04:25 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:04:25 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.313, Max reward: 0.940 +2025-04-04 at 04:04:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:04:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "oldest recorded historical data" +2025-04-04 at 04:04:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-04 at 04:04:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "when did time recording begin" +2025-04-04 at 04:04:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ +Result 2: +Time, hr:min Range station MILA Bandpass plots or tabs Bilevels Computers words O'graph records isruslh recrds :ipecial plots or tabs Speeial suo.tooal From To 00:00 00:02 00:04 00:07 00:14 01:31 01:33 01:48 02:25 02:34 02:43 02:49 04:44 08:35 12:49 13:18 16:44 17 :15 20 :37 24:53 27:01 37:33 40:55 44:39 00:12 00:14 04:44 00:18 00:28 01:33 01:45 01:59 02:34 02:45 02:50 13:59 08:35 12:49 16:44 17 :12 20 :37 25:00 BDA MSFN VAN CYI GDS MILA CYI CRO HAW HAW GDS MSFN MSFN MSFN HSK MSFN MAD 27:01 MSFN 37:42 GDS 40:55 MSFN 42:47 HSK 44:38 MS FN 52:37 MSFN 58:39 CDS 58:39 MS FN 101:58 GDS 123:12 GDS X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ + +2025-04-04 at 04:04:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the earliest recorded data point for a specific topic, possibly including its type, field or domain? +2025-04-04 at 04:04:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Time, hr:min Range station MILA Bandpass plots or tabs Bilevels Computers words O'graph records isruslh recrds :ipecial plots or tabs Speeial suo.tooal From To 00:00 00:02 00:04 00:07 00:14 01:31 01:33 01:48 02:25 02:34 02:43 02:49 04:44 08:35 12:49 13:18 16:44 17 :15 20 :37 24:53 27:01 37:33 40:55 44:39 00:12 00:14 04:44 00:18 00:28 01:33 01:45 01:59 02:34 02:45 02:50 13:59 08:35 12:49 16:44 17 :12 20 :37 25:00 BDA MSFN VAN CYI GDS MILA CYI CRO HAW HAW GDS MSFN MSFN MSFN HSK MSFN MAD 27:01 MSFN 37:42 GDS 40:55 MSFN 42:47 HSK 44:38 MS FN 52:37 MSFN 58:39 CDS 58:39 MS FN 101:58 GDS 123:12 GDS X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-04 at 04:04:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:04:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch data duration +2025-04-04 at 04:04:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-04 at 04:04:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Earliest recorded data point in history +2025-04-04 at 04:04:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-04 at 04:04:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: seismometers seismometer data Apollo 11 lunar launch perturbations electric field +2025-04-04 at 04:04:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-04 at 04:04:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:04:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission training timeline +2025-04-04 at 04:04:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:04:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: earliest seismic signal after Apollo 12 moon impact +2025-04-04 at 04:04:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:04:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Electric field measurements Apollo 12 launch perturbations atmospheric disturbances exhaust plume +2025-04-04 at 04:04:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:04:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:04:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 1969 Apollo mission training schedule +2025-04-04 at 04:04:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:04:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Farthest depth from which seismic signal consistency indicates rock material throughout lunar crust +2025-04-04 at 04:04:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-04 at 04:04:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 electrical disturbances launch power separation measurements +2025-04-04 at 04:04:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:04:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:04:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first data point in Apollo 11 mission archives +2025-04-04 at 04:04:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:04:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: density of moon material at different depths +2025-04-04 at 04:04:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ + +2025-04-04 at 04:04:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar module power system battery malfunction current surge +2025-04-04 at 04:04:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-04 at 04:04:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:04:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo seismic data record start time +2025-04-04 at 04:04:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:04:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moon internal material density and atmospheric structure +2025-04-04 at 04:04:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-04 at 04:04:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar module descent electrical system short circuit current transient +2025-04-04 at 04:04:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:04:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:04:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moon composition implications seismic data +2025-04-04 at 04:04:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:04:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-IVB impact site details +2025-04-04 at 04:04:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-04 at 04:04:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:04:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission objectives +2025-04-04 at 04:04:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:04:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-IVB impact data +2025-04-04 at 04:04:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:04:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:04:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew training schedule +2025-04-04 at 04:04:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:04:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB impact signal characteristics and implications for moon structure +2025-04-04 at 04:04:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:04:52 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:04:52 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:04:52 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1764, 1826, 1050, 482, 724, 487] +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_correctness:84 - Average student length: 1055.50 +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 5.00 +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_correctness:86 - Length ratio: 211.10 +2025-04-04 at 04:04:52 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-04 at 04:04:52 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.158 Âą 0.354 +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 1.00 Âą 2.24 +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 6, 0, 0, 0] +2025-04-04 at 04:04:52 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:04:52 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nBecause of access restrictions to sites 8 and 9, the corresponding recorders ...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nThe character of the signal from the S-IVB impact is identical to that of the...', 'Result 1:\na. Perform selenological inspection, survey, and sampling of materials in a p...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ +Result 2: +Time, hr:min Range station MILA Bandpass plots or tabs Bilevels Computers words O'graph records isruslh recrds :ipecial plots or tabs Speeial suo.tooal From To 00:00 00:02 00:04 00:07 00:14 01:31 01:33 01:48 02:25 02:34 02:43 02:49 04:44 08:35 12:49 13:18 16:44 17 :15 20 :37 24:53 27:01 37:33 40:55 44:39 00:12 00:14 04:44 00:18 00:28 01:33 01:45 01:59 02:34 02:45 02:50 13:59 08:35 12:49 16:44 17 :12 20 :37 25:00 BDA MSFN VAN CYI GDS MILA CYI CRO HAW HAW GDS MSFN MSFN MSFN HSK MSFN MAD 27:01 MSFN 37:42 GDS 40:55 MSFN 42:47 HSK 44:38 MS FN 52:37 MSFN 58:39 CDS 58:39 MS FN 101:58 GDS 123:12 GDS X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Time, hr:min Range station MILA Bandpass plots or tabs Bilevels Computers words O'graph records isruslh recrds :ipecial plots or tabs Speeial suo.tooal From To 00:00 00:02 00:04 00:07 00:14 01:31 01:33 01:48 02:25 02:34 02:43 02:49 04:44 08:35 12:49 13:18 16:44 17 :15 20 :37 24:53 27:01 37:33 40:55 44:39 00:12 00:14 04:44 00:18 00:28 01:33 01:45 01:59 02:34 02:45 02:50 13:59 08:35 12:49 16:44 17 :12 20 :37 25:00 BDA MSFN VAN CYI GDS MILA CYI CRO HAW HAW GDS MSFN MSFN MSFN HSK MSFN MAD 27:01 MSFN 37:42 GDS 40:55 MSFN 42:47 HSK 44:38 MS FN 52:37 MSFN 58:39 CDS 58:39 MS FN 101:58 GDS 123:12 GDS X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:04:52 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nTime, hr:min Range station MILA Bandpass plots or tabs Bilevels Computers wor...', 'Result 1:\nBecause of access restrictions to sites 8 and 9, the corresponding recorders ...', 'Result 1:\nAs a result of the electrical disturbances experienced during the Apollo l2 l...', 'Result 1:\nThe field-change and sferics detectors at site 5 gave no indication of any li...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nAt 97 hours 14 minutes, the crew reported a thumping noise and snowflakes ven...'] +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-04 at 04:04:52 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-04 at 04:04:52 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-04 at 04:04:52 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 1.0 +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.167 +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:04:52 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.367, Perfect scores: 1/6 +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.50 +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:04:52 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:04:52 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.336, Max reward: 0.782 +2025-04-04 at 04:04:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:04:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Propulsion auxiliary propellant gaging system failure impact on mission performance +2025-04-04 at 04:04:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:04:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "SMAL thruster failure impact on DOD space mission performance" +2025-04-04 at 04:04:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:04:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: service propulsion auxiliary propellant gaging system mission failure impact +2025-04-04 at 04:04:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:04:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what impact did the failure of the Mars 2020 Perseverance rover's auxiliary propellant gaging system have on the rover's overall mission performance? +2025-04-04 at 04:04:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:04:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission propulsion failure investigation +2025-04-04 at 04:04:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:04:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:05:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Effect of propellant leakage and venting on orbital maneuvering and navigation +2025-04-04 at 04:05:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ + +2025-04-04 at 04:05:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: STSDH-MSG-019 text "SPAGS failure" MSC-02680 +2025-04-04 at 04:05:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-04 at 04:05:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: implications of system failure propellant gaging system on mission performance +2025-04-04 at 04:05:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:05:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what performance impacts was the auxiliary propellant gaging system failure have on the Mars 2020 Propulsion Module's performance? +2025-04-04 at 04:05:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:05:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:05:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Spacecraft autopilot response to system failures with re- translation maneuver +2025-04-04 at 04:05:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-04 at 04:05:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: SPAGS failure DID not affect flight control systems +2025-04-04 at 04:05:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:05:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ë‹Ŧ propellant consumption reaction control system propellant venting +2025-04-04 at 04:05:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:05:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: did the failure of the auxiliary propellant gaging system before launch affect the Mars 2020 Perseverance rover's overall mission performance? +2025-04-04 at 04:05:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:05:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:05:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Spacecraft attitude calculation bias during digital autopilot initialization +2025-04-04 at 04:05:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-04 at 04:05:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 SPAGS failure cause +2025-04-04 at 04:05:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:05:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system propellant usage lunar landing +2025-04-04 at 04:05:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:05:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: did the primary propellant gaging system of the Mars 2020 Perseverance rover perform as expected? +2025-04-04 at 04:05:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:05:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:05:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Attitude error bias during sixth midcourse correction +2025-04-04 at 04:05:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +initial maneuver to the firing attitude for the final midcourse correction was done manually using the earth as a reference in the same manner as the previous maneuver. This procedure presented no problems , even though the earth disk was considerably larger at this time. +------ + +2025-04-04 at 04:05:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system abort handling lunar landing +2025-04-04 at 04:05:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:05:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: were the primary gaging system and reaction control system's propellant consumption within nominal limits for the Mars 2020 Perseverance rover mission? +2025-04-04 at 04:05:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:05:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:05:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Manual attitude maneuvers and attitude error bias +2025-04-04 at 04:05:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-04 at 04:05:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar module pitchdown 10 degrees after reaction control system failure +2025-04-04 at 04:05:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:05:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: did the primary propellant gaging system of the Mars 2020 Perseverance rover operate within expected parameters during the mission? +2025-04-04 at 04:05:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:05:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:05:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Digital autopilot initialization errors as related to attitude bias +2025-04-04 at 04:05:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-04 at 04:05:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch maneuver and control after reaction control failure lunar landing +2025-04-04 at 04:05:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:05:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: did the primary propellant gaging system of the Mars 2020 Perseverance rover perform nominally during the mission? +2025-04-04 at 04:05:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:05:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:05:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Digital autopilot initialization and navigation system integrity +2025-04-04 at 04:05:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-04 at 04:05:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitchdown maneuvering rules Apollo +2025-04-04 at 04:05:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ + +2025-04-04 at 04:05:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:05:21 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:05:21 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:05:22 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, True, False, False, False] +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_correctness:82 - Student lengths: [2002, 464, 681, 1565, 424, 717] +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [2, 2, 2, 2, 2, 2] +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_correctness:84 - Average student length: 975.50 +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 2.00 +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_correctness:86 - Length ratio: 487.75 +2025-04-04 at 04:05:22 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:05:22 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.371 Âą 0.396 +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 4.67 Âą 5.85 +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [16, 4, 8, 0, 0, 0] +2025-04-04 at 04:05:22 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +initial maneuver to the firing attitude for the final midcourse correction was done manually using the earth as a reference in the same manner as the previous maneuver. This procedure presented no problems , even though the earth disk was considerably larger at this time. +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:05:22 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-04 at 04:05:22 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:05:22 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.333, Perfect scores: 2/6 +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 2/6 +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.50 +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:05:22 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:05:22 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.450, Max reward: 0.940 +2025-04-04 at 04:05:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:05:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericynthion altitude translunar injection +2025-04-04 at 04:05:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-04 at 04:05:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericynthion altitude apollo 11 translunar injection +2025-04-04 at 04:05:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:05:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection altitude +2025-04-04 at 04:05:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-04 at 04:05:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericynthion altitude translunar injection +2025-04-04 at 04:05:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-04 at 04:05:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: `pericynthion altitude translunar injection` +2025-04-04 at 04:05:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-04 at 04:05:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo translunar inclination and pericynthion altitude +2025-04-04 at 04:05:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:05:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:05:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericynthion altitude translunar injection apollo 11 +2025-04-04 at 04:05:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:05:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericynthion altitude at apollo 11 TLI +2025-04-04 at 04:05:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:05:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection altitude +2025-04-04 at 04:05:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-04 at 04:05:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: translunar injection Apollo 11 S-IVB +2025-04-04 at 04:05:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:05:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:05:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericynthion altitude translunar to free return +2025-04-04 at 04:05:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:05:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: perfected pericynthion altitude translunar injection +2025-04-04 at 04:05:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-04 at 04:05:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar phasing transearth injection +2025-04-04 at 04:05:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:05:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: translunar injection Apollo 11 S-IVB +2025-04-04 at 04:05:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:05:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:05:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 translunar injection pericynthion altitude +2025-04-04 at 04:05:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:05:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection timing +2025-04-04 at 04:05:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:05:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: translunar injection pericynthion altitude Apollo 11 +2025-04-04 at 04:05:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:05:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:05:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo mission synopses +2025-04-04 at 04:05:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:05:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection timing in chronological order +2025-04-04 at 04:05:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:05:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module translunar injection maneuver Apollo 11 +2025-04-04 at 04:05:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:05:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:05:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 10 transearth injection +2025-04-04 at 04:05:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:05:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 mission sequence table numbers +2025-04-04 at 04:05:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:05:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module translunar injection altitude +2025-04-04 at 04:05:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:05:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:05:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 10 transearth injection pericynthion altitude +2025-04-04 at 04:05:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:05:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 supplement 3 lunar module abort guidance system performance +2025-04-04 at 04:05:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:05:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:05:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 10 s-ivb transearth injection pericynthion altitude +2025-04-04 at 04:05:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:05:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 1 abort profile translunar injection altitude +2025-04-04 at 04:05:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-04 at 04:05:51 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:05:51 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:05:51 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, True, False, True] +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1914, 479, 1689, 793, 281, 312] +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [11, 11, 11, 11, 11, 11] +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_correctness:84 - Average student length: 911.33 +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 11.00 +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_correctness:86 - Length ratio: 82.85 +2025-04-04 at 04:05:51 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-04 at 04:05:51 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.296 Âą 0.348 +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 3.67 Âą 5.50 +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [15, 0, 0, 6, 0, 1] +2025-04-04 at 04:05:51 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-04 at 04:05:51 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\n8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH CO...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...'] +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 5.0 +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.833 +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:05:51 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.267, Perfect scores: 1/6 +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 2/6 +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.67 +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:05:51 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:05:51 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.220, Max reward: 0.483 +2025-04-04 at 04:05:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:05:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Command Module moment of inertia +2025-04-04 at 04:05:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:05:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command and service module moment of inertia +2025-04-04 at 04:05:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-04 at 04:05:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command and Service Module moment of inertia +2025-04-04 at 04:05:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-04 at 04:05:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the moment of inertia of the Command and Service Module on the International Space Station +2025-04-04 at 04:05:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-04 at 04:05:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:05:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command and Service Module physical dynamics moment of inertia +2025-04-04 at 04:05:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:05:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command and service module CSM moment of inertia values +2025-04-04 at 04:05:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:05:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment of inertia of a Command & Service Module of the Space Shuttle +2025-04-04 at 04:05:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-04 at 04:05:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: dimension of the Command and Service Module +2025-04-04 at 04:05:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-04 at 04:05:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:06:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Command and Service Module mass and moment of inertia +2025-04-04 at 04:06:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:06:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment of inertia reference CSM +2025-04-04 at 04:06:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:06:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: SpaceX Starlink Command and Service Module physical parameters +2025-04-04 at 04:06:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:06:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: dimensions of the Command Service Module +2025-04-04 at 04:06:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-04 at 04:06:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:06:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Command and Service Module mass +2025-04-04 at 04:06:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:06:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command and service module NASA specifications +2025-04-04 at 04:06:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:06:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Starlink Command and Service Module moment of inertia +2025-04-04 at 04:06:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:06:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: dimensions Command and Service Module +2025-04-04 at 04:06:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-04 at 04:06:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:06:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Command and Service Module inertia values +2025-04-04 at 04:06:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-04 at 04:06:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: technical documentation for Apollo CSM inertia +2025-04-04 at 04:06:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-04 at 04:06:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment of inertia of Starlink or SpaceX Command and Service Module +2025-04-04 at 04:06:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-04 at 04:06:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:06:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Calculate translational inertia from velocity vector +2025-04-04 at 04:06:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-04 at 04:06:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA CSM physics reference materials +2025-04-04 at 04:06:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-04 at 04:06:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: SpaceX Starlink Command and Service Module technical specifications +2025-04-04 at 04:06:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:06:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:06:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Calculate translational inertia from velocity vector and acceleration +2025-04-04 at 04:06:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:06:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA document MSC-02680 +2025-04-04 at 04:06:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:06:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Stowable Starlink Command and Service Module mass +2025-04-04 at 04:06:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:06:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:06:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo CSM official documents +2025-04-04 at 04:06:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:06:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Starlink Command and Service Module physical dimensions +2025-04-04 at 04:06:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-04 at 04:06:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:06:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Apollo CSM mass properties +2025-04-04 at 04:06:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:06:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Small satellite moment of inertia formulas +2025-04-04 at 04:06:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:06:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:06:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: SpaceX Command and Service Module mass and moment of inertia +2025-04-04 at 04:06:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-04 at 04:06:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:06:20 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:06:20 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:06:20 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_correctness:82 - Student lengths: [567, 314, 1011, 561, 385, 879] +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_correctness:84 - Average student length: 619.50 +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 5.00 +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_correctness:86 - Length ratio: 123.90 +2025-04-04 at 04:06:20 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_format:228 - Responses ending properly: 0/6 +2025-04-04 at 04:06:20 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.000 Âą 0.000 +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 0.00 Âą 0.00 +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 0, 0, 0, 0] +2025-04-04 at 04:06:20 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-04 at 04:06:20 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-04 at 04:06:20 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:06:20 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 1/6 +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.00 +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 0/6 +2025-04-04 at 04:06:20 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:06:20 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.212, Max reward: 0.603 +2025-04-04 at 04:06:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:06:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module Pilot antenna adjustments Apollo 8 +2025-04-04 at 04:06:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ + +2025-04-04 at 04:06:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 command module antenna angle adjustment" +2025-04-04 at 04:06:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:06:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 antenna settings adjustment Commander +2025-04-04 at 04:06:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:06:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo mission antenna angle adjustment +2025-04-04 at 04:06:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ + +2025-04-04 at 04:06:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Command Module Pilot antenna settings pitch adjustment +2025-04-04 at 04:06:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:06:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module pilot antenna pitch adjustment Apollo 11 +2025-04-04 at 04:06:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:06:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:06:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 crew command module spaceship antenna pitch adjustments +2025-04-04 at 04:06:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:06:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"lunar module pitch angle adjustment Apollo 11" +2025-04-04 at 04:06:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:06:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 communications shunt switch +2025-04-04 at 04:06:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:06:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo lunar descent angle adjustment crew command module pilot +2025-04-04 at 04:06:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:06:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Command Module Pilot manual pitch maneuver during lunar module docking +2025-04-04 at 04:06:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-04 at 04:06:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitching angle for manual antenna setting Apollo 11 +2025-04-04 at 04:06:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:06:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:06:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 pitch angle command module +2025-04-04 at 04:06:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:06:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +" Transearth injection maneuver pitch angle adjustment Apollo 11" +2025-04-04 at 04:06:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:06:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 circuit breaker position +2025-04-04 at 04:06:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-04 at 04:06:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Swinging the lunar module command module antenna to position the sun +2025-04-04 at 04:06:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +Lunar module powerup for the final midcourse correction maneuver was performed according to the prescribed contingency checklist, with only minor deviations furnished by the ground. Shortly afterward, the lunar module windows cleared of moisture and the cabin temperature again became comfortable. Approximately 6 hours before entry, the passive thermal control mode vwas terminated and the spacecraft was maneuvered to place the earth in the crewmen optical alignment sight with the terminator parallel to the Y axis in preparation for the midcourse maneuver. At that time, a sun/moon alignment was made. Acquisition of these bodies was made by pitching up in a plane roughly parallel to the ecliptic plane. The sun filter made viewing through the telescope reticle very difficult. The spacecraft was controlled by the Lunar Module Pilot from commands given by the Commander, who responded when the reticle lines bisected the moon and solar disks. Three sets of marks were taken on each body. The +------ + +2025-04-04 at 04:06:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Pitch control rate of 1.5 degrees per second +2025-04-04 at 04:06:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:06:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 attitude control pitch maneuver lunar module +2025-04-04 at 04:06:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:06:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:06:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"lunar mission post abort sudden inertial velocity Apollo 11" +2025-04-04 at 04:06:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:06:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: l0 percent throttle pitch adjustment +2025-04-04 at 04:06:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ + +2025-04-04 at 04:06:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module sun/moon alignment angle +2025-04-04 at 04:06:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-04 at 04:06:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 descent pitch maneuver +2025-04-04 at 04:06:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:06:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:06:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"S-IVB impact coordinates and velocity after first transearth correction Apollo 11" +2025-04-04 at 04:06:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:06:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module antenna pitch maximum excursion +2025-04-04 at 04:06:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:06:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module platform alignment at entrance into lunar orbit using the sun +2025-04-04 at 04:06:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-04 at 04:06:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Transearth injection maneuver Apollo 11 +2025-04-04 at 04:06:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-04 at 04:06:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:06:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Lunar S-IVB impact signal characteristics and interpretation" +2025-04-04 at 04:06:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:06:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module pitch maneuver 2lo-foot receiving antenna +2025-04-04 at 04:06:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:06:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 midcourse correction +2025-04-04 at 04:06:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:06:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:06:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"why the S-IVB signal shows a rapid buildup during impulse" +2025-04-04 at 04:06:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:06:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module undocking +2025-04-04 at 04:06:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:06:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:06:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: + "Seismic signal buildup and acceleration during S-IVB impact" +2025-04-04 at 04:06:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:06:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 undocking procedure +2025-04-04 at 04:06:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:06:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:06:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Lunar impact similarity to natural meteoroid signals" +2025-04-04 at 04:06:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:06:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 undocking and separation procedures +2025-04-04 at 04:06:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:06:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:06:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"S-IVB impact signal velocity to l00 km depth" +2025-04-04 at 04:06:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:06:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 re-entry procedures +2025-04-04 at 04:06:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:06:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:06:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: altitude at Apollo 11 landing +2025-04-04 at 04:06:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:06:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:06:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 landing latitude +2025-04-04 at 04:06:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:06:45 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:06:45 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:06:45 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, True, False, False, False] +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_correctness:82 - Student lengths: [132, 1809, 642, 389, 231, 2015] +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_correctness:84 - Average student length: 869.67 +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_correctness:86 - Length ratio: 86.97 +2025-04-04 at 04:06:45 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-04 at 04:06:45 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.188 Âą 0.270 +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 1.67 Âą 2.62 +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [3, 0, 7, 0, 0, 0] +2025-04-04 at 04:06:45 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:06:45 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ["Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...'] +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:06:45 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe lunar module platform was coarse aligned to the command module platform a...', 'Result 1:\nThe lunar module platform was coarse aligned to the command module platform a...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe discarded S-IVB stage was targeted for a lunar impact of 3 degrees south ...', 'Result 1:\nThe discarded S-IVB stage was targeted for a lunar impact of 3 degrees south ...', 'Result 1:\ncrewmen noted the small change in acceleration caused by the mixture ratio sh...', 'Result 1:\nAn unexplained characteristic of the S-IVB impact is the rapid buildup from i...', 'Result 1:\nThe character of the signal from the S-IVB impact is identical to that of the...', 'Result 1:\nAn unexplained characteristic of the S-IVB impact is the rapid buildup from i...'] +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:06:45 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\n6.3 COMMUNICATIONS EQUIPMENT\n\nS-band communications were nominal from system ...', 'Result 1:\n6.3 COMMUNICATIONS EQUIPMENT\n\nS-band communications were nominal from system ...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nCondition Maneuver Second midcourse correction Transearth injection Third mid...', 'Result 1:\n6.3 COMMUNICATIONS EQUIPMENT\n\nS-band communications were nominal from system ...', 'Result 1:\n6.3 COMMUNICATIONS EQUIPMENT\n\nS-band communications were nominal from system ...'] +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +Lunar module powerup for the final midcourse correction maneuver was performed according to the prescribed contingency checklist, with only minor deviations furnished by the ground. Shortly afterward, the lunar module windows cleared of moisture and the cabin temperature again became comfortable. Approximately 6 hours before entry, the passive thermal control mode vwas terminated and the spacecraft was maneuvered to place the earth in the crewmen optical alignment sight with the terminator parallel to the Y axis in preparation for the midcourse maneuver. At that time, a sun/moon alignment was made. Acquisition of these bodies was made by pitching up in a plane roughly parallel to the ecliptic plane. The sun filter made viewing through the telescope reticle very difficult. The spacecraft was controlled by the Lunar Module Pilot from commands given by the Commander, who responded when the reticle lines bisected the moon and solar disks. Three sets of marks were taken on each body. The +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-04 at 04:06:45 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nAt approximately 105 hours, the crew performed a manual descent propulsion ma...', 'Result 1:\nAt approximately 105 hours, the crew performed a manual descent propulsion ma...', 'Result 1:\nAt 2-l/2 hours prior to entry, the command module was fully powered up and lu...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...'] +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:06:45 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...'] +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 12: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:06:45 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ["Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...', 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nIn this report, all actual times prior to earth landing are elapsed time from...', 'Result 1:\nIn this report, all actual times prior to earth landing are elapsed time from...'] +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 0.0 +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.000 +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.000 +2025-04-04 at 04:06:45 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.400, Perfect scores: 0/6 +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.67 +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:06:45 | INFO | src.rewards:log_chat_state:855 - īŋŊīŋŊīŋŊ Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:06:45 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.545, Max reward: 0.720 +2025-04-04 at 04:06:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:06:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission emergency abort +2025-04-04 at 04:06:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-04 at 04:06:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 emergency abort swastika capsule +2025-04-04 at 04:06:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:06:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 emergency abort +2025-04-04 at 04:06:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:06:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission +2025-04-04 at 04:06:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:06:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo mission with emergency abort +2025-04-04 at 04:06:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-04 at 04:06:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:06:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 oxygen tank fire +2025-04-04 at 04:06:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:06:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 explosion details +2025-04-04 at 04:06:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-04 at 04:06:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission abort reasons +2025-04-04 at 04:06:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:06:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: CO2 scrubbing in space exploration +2025-04-04 at 04:06:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:06:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:06:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Reasons Apollo 13 first mission to require emergency abort +2025-04-04 at 04:06:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +4.0 TRAJECTORY + +The planned trajectory profile was similar to that for Apollo 12 except for descent orbit insertion being performed with the service propulsion system and the targeting of the spent S--IvB stage for a lunar impact. The trajectory had been very close to the nominal flight plan up to the time of abort, which was the first in the Apollo program. Throughout the manned space program, techniques have been developed and tested for the real-tine determination of immediate abort requirements, but Apollo l3 presented the first situation in which their use was necesSary. Figure 3-l shows the mission profile, including the relative locations of all major maneuvers. +------ + +2025-04-04 at 04:06:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 explosion oxygen tank failure +2025-04-04 at 04:06:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:06:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 explosion cause +2025-04-04 at 04:06:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-04 at 04:06:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: carbon dioxide scrubbing in lunar module +2025-04-04 at 04:06:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-04 at 04:06:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:07:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 system malfunctions and scientific challenges +2025-04-04 at 04:07:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:07:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 explosion oxygen tank failure afterlava +2025-04-04 at 04:07:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-04 at 04:07:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar impact +2025-04-04 at 04:07:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:07:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module operations during transearth coast +2025-04-04 at 04:07:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:07:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:07:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 cryogenic oxygen supply anomaly analysis +2025-04-04 at 04:07:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:07:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 oxygen tank 1 failure sequence +2025-04-04 at 04:07:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:07:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 training geology +2025-04-04 at 04:07:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:07:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module re-entry propulsion +2025-04-04 at 04:07:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:07:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:07:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module cryogenic oxygen system operations +2025-04-04 at 04:07:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:07:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 houston mission control summary +2025-04-04 at 04:07:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:07:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission reports +2025-04-04 at 04:07:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:07:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module descent engine firing +2025-04-04 at 04:07:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:07:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:07:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lifeboat mode cryogenic oxygen system +2025-04-04 at 04:07:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:07:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 crew training requirements +2025-04-04 at 04:07:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:07:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission reports +2025-04-04 at 04:07:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:07:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module emergency activation +2025-04-04 at 04:07:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:07:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:07:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lifeboat mode cryogenic oxygen system operations details +2025-04-04 at 04:07:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:07:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 crew training geology experiment +2025-04-04 at 04:07:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-04 at 04:07:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission analysis +2025-04-04 at 04:07:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:07:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 command module emergency powering sequence +2025-04-04 at 04:07:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:07:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:07:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar landing training objectives +2025-04-04 at 04:07:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-04 at 04:07:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module electrical power system +2025-04-04 at 04:07:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:07:16 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:07:16 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:07:16 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, True, False] +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_correctness:82 - Student lengths: [863, 1758, 1956, 1983, 346, 394] +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [9, 9, 9, 9, 9, 9] +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_correctness:84 - Average student length: 1216.67 +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 9.00 +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_correctness:86 - Length ratio: 135.19 +2025-04-04 at 04:07:16 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:07:16 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.296 Âą 0.348 +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 4.50 Âą 6.68 +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [8, 0, 0, 18, 1, 0] +2025-04-04 at 04:07:16 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +4.0 TRAJECTORY + +The planned trajectory profile was similar to that for Apollo 12 except for descent orbit insertion being performed with the service propulsion system and the targeting of the spent S--IvB stage for a lunar impact. The trajectory had been very close to the nominal flight plan up to the time of abort, which was the first in the Apollo program. Throughout the manned space program, techniques have been developed and tested for the real-tine determination of immediate abort requirements, but Apollo l3 presented the first situation in which their use was necesSary. Figure 3-l shows the mission profile, including the relative locations of all major maneuvers. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:07:16 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nin figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit ci...', 'Result 1:\nadvised of their consumables status. A procedure was developed on the ground ...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe unprecedented powered-down state of the command module required generatio...', 'Result 1:\nThe unprecedented powered-down state of the command module required generatio...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...'] +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-04 at 04:07:16 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:07:16 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.533, Perfect scores: 2/6 +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 2.17 +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 3/6 +2025-04-04 at 04:07:16 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:07:16 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.345, Max reward: 0.601 +2025-04-04 at 04:07:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:07:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: spongebob squarepants entry batteries +2025-04-04 at 04:07:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ +Result 2: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ + +2025-04-04 at 04:07:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "original battery capacity Tesla Roadster +2025-04-04 at 04:07:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-04 at 04:07:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Interstellar Christopher Nolan entry batteries energy" +2025-04-04 at 04:07:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-04 at 04:07:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:07:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 mission battery capacity and energy consumption" +2025-04-04 at 04:07:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:07:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:07:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 mission entry battery energy remaining" +2025-04-04 at 04:07:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-04 at 04:07:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:07:25 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:07:25 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:07:25 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, True, False, False] +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_correctness:82 - Student lengths: [35, 483, 370, 295, 191, 343] +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [15, 15, 15, 15, 15, 15] +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_correctness:84 - Average student length: 286.17 +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 15.00 +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_correctness:86 - Length ratio: 19.08 +2025-04-04 at 04:07:25 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-04 at 04:07:25 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.167 Âą 0.251 +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 0.67 Âą 1.11 +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 3, 1, 0, 0] +2025-04-04 at 04:07:25 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-04 at 04:07:25 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ +Result 2: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ + +2025-04-04 at 04:07:25 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nCommand module battery performance was acceptable throughout the mission. Ent...'] +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-04 at 04:07:25 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...'] +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-04 at 04:07:25 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe command module was completely powered down at 58 hours 40 minutes , at wh...'] +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-04 at 04:07:25 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-04 at 04:07:25 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 0.0 +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.000 +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.000 +2025-04-04 at 04:07:25 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.167, Perfect scores: 1/6 +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 1/6 +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 0.33 +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:07:25 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:07:25 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.176, Max reward: 0.653 +2025-04-04 at 04:07:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:07:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What biomedical signals monitored by the Apollo Lunar Module? +2025-04-04 at 04:07:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:07:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how many biomedical signals can be monitored in the lunar module +2025-04-04 at 04:07:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-04 at 04:07:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA Apollo biomedical data channels" +2025-04-04 at 04:07:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:07:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Biomedical signals monitored in Lunar Module instrumentation +2025-04-04 at 04:07:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:07:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how many channels are monitored in the Apollo Lunar Module's System +2025-04-04 at 04:07:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:07:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How many biomedical monitors were used in the Apollo Lunar Module +2025-04-04 at 04:07:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:07:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:07:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Maximum biomedical signals monitorable on Lunar Module +2025-04-04 at 04:07:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:07:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module biomedical signals simultaneous monitoring +2025-04-04 at 04:07:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:07:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo Lunar Module medical data capabilities" +2025-04-04 at 04:07:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-04 at 04:07:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Biomedical monitoring system Lunar Module Capabilities +2025-04-04 at 04:07:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-04 at 04:07:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: biomedical signals monitored in Apollo Lunar Module +2025-04-04 at 04:07:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:07:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How many biomedical monitors could astronauts have in the Lunar Module's command or lunar module +2025-04-04 at 04:07:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-04 at 04:07:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:07:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Can the Lunar Module monitor multiple onboard crew members simultaneously? +2025-04-04 at 04:07:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-04 at 04:07:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo lunar module biomedical signal monitoring capabilities +2025-04-04 at 04:07:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:07:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo Lunar Module monitoring limitations" +2025-04-04 at 04:07:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:07:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Biomedical monitoring system multi-channel capabilities +2025-04-04 at 04:07:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-04 at 04:07:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Lunar Module limit biomedical signals +2025-04-04 at 04:07:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The major medical concern, recogmized immediately after the abort decision, was the possibility of carbon dioxide buildup in the lunar module atmosphere. Since the physiological effects of increased carbon dioxide concentration are well known and readily recognizable with proper biomedical monitoring, the allowable limit of carbon dioxide buildup was increased from the nominal 7.6 to 15mm Hg. The carbon dioxide level was above 7.6mm Hg for only a 4-hour period, and no adverse physiological effects or degradation in crew performance resulted from this elevated concentration. Modified use of the lithium hydroxide cartridges (section 6.7) maintained the carbon dioxide partial pressure well below lmm Hg for the remainder of the flight. + +9.2.3 Sleep +------ + +2025-04-04 at 04:07:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Lunar Module design limitations on simultaneous biomedical monitoring +2025-04-04 at 04:07:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:07:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:07:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Can Lunar Module S-band communications handle multiple signals? +2025-04-04 at 04:07:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:07:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: single-person electrocardiogram signal monitoring +2025-04-04 at 04:07:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-04 at 04:07:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 lunar module systems" +2025-04-04 at 04:07:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:07:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Biomedical monitoring system for a 3-person crew in Lunar Module +2025-04-04 at 04:07:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-04 at 04:07:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Lunar Module onboard crew monitoring +2025-04-04 at 04:07:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:07:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Lunar Module physiology experiment unrelated to biological monitoring +2025-04-04 at 04:07:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:07:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:07:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-band communication equipment multiple channel capability +2025-04-04 at 04:07:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:07:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module single crew member electrocardiogram monitoring +2025-04-04 at 04:07:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-04 at 04:07:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 Lunar Module biomedical monitoring limitations" +2025-04-04 at 04:07:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The major medical concern, recogmized immediately after the abort decision, was the possibility of carbon dioxide buildup in the lunar module atmosphere. Since the physiological effects of increased carbon dioxide concentration are well known and readily recognizable with proper biomedical monitoring, the allowable limit of carbon dioxide buildup was increased from the nominal 7.6 to 15mm Hg. The carbon dioxide level was above 7.6mm Hg for only a 4-hour period, and no adverse physiological effects or degradation in crew performance resulted from this elevated concentration. Modified use of the lithium hydroxide cartridges (section 6.7) maintained the carbon dioxide partial pressure well below lmm Hg for the remainder of the flight. + +9.2.3 Sleep +------ + +2025-04-04 at 04:07:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Biomedical monitoring system crew size limitations +2025-04-04 at 04:07:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-04 at 04:07:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 mission crew biomedical signals +2025-04-04 at 04:07:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-04 at 04:07:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo biomedical telemetry and monitoring protocols +2025-04-04 at 04:07:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:07:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:07:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-band transmission equipment theoretical bandwidth +2025-04-04 at 04:07:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:07:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar module carbon dioxide buildup Apollo 13" +2025-04-04 at 04:07:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:07:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Biomedical monitoring system entry phase simultaneous +2025-04-04 at 04:07:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-04 at 04:07:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: heart rate data Lunar Module Pilot +2025-04-04 at 04:07:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-04 at 04:07:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo communication systems for S-band transmissions and power management +2025-04-04 at 04:07:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-04 at 04:07:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:07:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 carbon dioxide removal systems" +2025-04-04 at 04:07:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:07:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Biomedical monitoring system crew operations phases +2025-04-04 at 04:07:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-04 at 04:07:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission telemetry systems defectivity +2025-04-04 at 04:07:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-04 at 04:07:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:07:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Command module lithium hydroxide cartridges biomedical applications" +2025-04-04 at 04:07:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The installed primary lithium hydroxide cartridge was used for approximately 27 hours (82 man-hoursīŧ‰ following activation of the lunar moduie at about 58 hours. The secondary cartridge was selected at about $\upbeta5$ 1/2 hours. During operation of the lunar module carbon dioxide removal system, the level was permitted to increase to an indicated l4.9 mm Hg. The primary cartridge is nominally rated for a usage capacity of 4l manhours at 520 Btu/man-hour. The secondary cartridge, nominally rated for about 17.9 man-hours, was used for 8 1/2 hours (25 1/2 man-hours). This cartridge is identical to that used in the portable life support system. A second primary cartridge was installed and used for approximately 6 minutes, but for the remainder of the mission, command module lithium hyaroxide cartridges were operated in a special arrangement. One side of each of two command module cartridges was covered and sealed with a plastic bag normally used to store a liquid-cooling garment. As shown +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:07:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: biomedical monitoring system ingress egress timelines +2025-04-04 at 04:07:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-04 at 04:07:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission recovery and signal loss +2025-04-04 at 04:07:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:07:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:07:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lithium Hydroxide cartridge lifespan in reducing carbon dioxide levels" +2025-04-04 at 04:07:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The installed primary lithium hydroxide cartridge was used for approximately 27 hours (82 man-hoursīŧ‰ following activation of the lunar moduie at about 58 hours. The secondary cartridge was selected at about $\upbeta5$ 1/2 hours. During operation of the lunar module carbon dioxide removal system, the level was permitted to increase to an indicated l4.9 mm Hg. The primary cartridge is nominally rated for a usage capacity of 4l manhours at 520 Btu/man-hour. The secondary cartridge, nominally rated for about 17.9 man-hours, was used for 8 1/2 hours (25 1/2 man-hours). This cartridge is identical to that used in the portable life support system. A second primary cartridge was installed and used for approximately 6 minutes, but for the remainder of the mission, command module lithium hyaroxide cartridges were operated in a special arrangement. One side of each of two command module cartridges was covered and sealed with a plastic bag normally used to store a liquid-cooling garment. As shown +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-04 at 04:07:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Biomedical monitoring system egress phase +2025-04-04 at 04:07:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-04 at 04:07:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Mission Data Losses System Complexity and Monitoring +2025-04-04 at 04:07:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-04 at 04:07:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:07:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: biomedical monitoring system egress procedure +2025-04-04 at 04:07:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-04 at 04:07:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Mission abort procedures and crew safety monitoring +2025-04-04 at 04:07:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-04 at 04:07:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:07:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Continuous monitoring systems after Apollo-style emergency abort +2025-04-04 at 04:07:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +4.0 TRAJECTORY + +The planned trajectory profile was similar to that for Apollo 12 except for descent orbit insertion being performed with the service propulsion system and the targeting of the spent S--IvB stage for a lunar impact. The trajectory had been very close to the nominal flight plan up to the time of abort, which was the first in the Apollo program. Throughout the manned space program, techniques have been developed and tested for the real-tine determination of immediate abort requirements, but Apollo l3 presented the first situation in which their use was necesSary. Figure 3-l shows the mission profile, including the relative locations of all major maneuvers. +------ + +2025-04-04 at 04:07:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:07:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo-style aborted lunar landing emergency medical monitoring +2025-04-04 at 04:07:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-04 at 04:07:55 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:07:55 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:07:55 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, False, True, True, False] +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.67 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_correctness:82 - Student lengths: [519, 18, 1915, 333, 179, 1541] +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [3, 3, 3, 3, 3, 3] +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_correctness:84 - Average student length: 750.83 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 3.00 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_correctness:86 - Length ratio: 250.28 +2025-04-04 at 04:07:55 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_format:228 - Responses ending properly: 6/6 +2025-04-04 at 04:07:55 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.871 Âą 0.177 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 9.67 Âą 4.85 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [6, 5, 19, 10, 6, 12] +2025-04-04 at 04:07:55 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The major medical concern, recogmized immediately after the abort decision, was the possibility of carbon dioxide buildup in the lunar module atmosphere. Since the physiological effects of increased carbon dioxide concentration are well known and readily recognizable with proper biomedical monitoring, the allowable limit of carbon dioxide buildup was increased from the nominal 7.6 to 15mm Hg. The carbon dioxide level was above 7.6mm Hg for only a 4-hour period, and no adverse physiological effects or degradation in crew performance resulted from this elevated concentration. Modified use of the lithium hydroxide cartridges (section 6.7) maintained the carbon dioxide partial pressure well below lmm Hg for the remainder of the flight. + +9.2.3 Sleep +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The installed primary lithium hydroxide cartridge was used for approximately 27 hours (82 man-hoursīŧ‰ following activation of the lunar moduie at about 58 hours. The secondary cartridge was selected at about $\upbeta5$ 1/2 hours. During operation of the lunar module carbon dioxide removal system, the level was permitted to increase to an indicated l4.9 mm Hg. The primary cartridge is nominally rated for a usage capacity of 4l manhours at 520 Btu/man-hour. The secondary cartridge, nominally rated for about 17.9 man-hours, was used for 8 1/2 hours (25 1/2 man-hours). This cartridge is identical to that used in the portable life support system. A second primary cartridge was installed and used for approximately 6 minutes, but for the remainder of the mission, command module lithium hyaroxide cartridges were operated in a special arrangement. One side of each of two command module cartridges was covered and sealed with a plastic bag normally used to store a liquid-cooling garment. As shown +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The installed primary lithium hydroxide cartridge was used for approximately 27 hours (82 man-hoursīŧ‰ following activation of the lunar moduie at about 58 hours. The secondary cartridge was selected at about $\upbeta5$ 1/2 hours. During operation of the lunar module carbon dioxide removal system, the level was permitted to increase to an indicated l4.9 mm Hg. The primary cartridge is nominally rated for a usage capacity of 4l manhours at 520 Btu/man-hour. The secondary cartridge, nominally rated for about 17.9 man-hours, was used for 8 1/2 hours (25 1/2 man-hours). This cartridge is identical to that used in the portable life support system. A second primary cartridge was installed and used for approximately 6 minutes, but for the remainder of the mission, command module lithium hyaroxide cartridges were operated in a special arrangement. One side of each of two command module cartridges was covered and sealed with a plastic bag normally used to store a liquid-cooling garment. As shown +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The major medical concern, recogmized immediately after the abort decision, was the possibility of carbon dioxide buildup in the lunar module atmosphere. Since the physiological effects of increased carbon dioxide concentration are well known and readily recognizable with proper biomedical monitoring, the allowable limit of carbon dioxide buildup was increased from the nominal 7.6 to 15mm Hg. The carbon dioxide level was above 7.6mm Hg for only a 4-hour period, and no adverse physiological effects or degradation in crew performance resulted from this elevated concentration. Modified use of the lithium hydroxide cartridges (section 6.7) maintained the carbon dioxide partial pressure well below lmm Hg for the remainder of the flight. + +9.2.3 Sleep +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +4.0 TRAJECTORY + +The planned trajectory profile was similar to that for Apollo 12 except for descent orbit insertion being performed with the service propulsion system and the targeting of the spent S--IvB stage for a lunar impact. The trajectory had been very close to the nominal flight plan up to the time of abort, which was the first in the Apollo program. Throughout the manned space program, techniques have been developed and tested for the real-tine determination of immediate abort requirements, but Apollo l3 presented the first situation in which their use was necesSary. Figure 3-l shows the mission profile, including the relative locations of all major maneuvers. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 12: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 6.0 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 1.000 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.000 +2025-04-04 at 04:07:55 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.600, Perfect scores: 3/6 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 4.17 +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 3/6 +2025-04-04 at 04:07:55 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:07:55 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.545, Max reward: 0.773 +2025-04-04 at 04:07:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:08:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 oxygen tank leak decision +2025-04-04 at 04:08:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-04 at 04:08:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 oxygen tank explosion mission decision +2025-04-04 at 04:08:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:08:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 1 cryogenic oxygen tank failure mission decision +2025-04-04 at 04:08:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-04 at 04:08:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Tanks containing cryogenic oxygen incident primary mission decision" +2025-04-04 at 04:08:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-04 at 04:08:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Oxygen tank incident military operation abort protocols 1940s" +2025-04-04 at 04:08:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ + +2025-04-04 at 04:08:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:08:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 oxygen leak management procedure +2025-04-04 at 04:08:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-04 at 04:08:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 hydrogen tank warning alarm decision +2025-04-04 at 04:08:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:08:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Apollo 11 crew decision after oxygen tank failure" +2025-04-04 at 04:08:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-04 at 04:08:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:08:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 oxygen leak repair alternatives +2025-04-04 at 04:08:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:08:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 electrical discharge thunderstorm northing +2025-04-04 at 04:08:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-04 at 04:08:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Abol 13 oxygen tank after Apollo 13 mission" +2025-04-04 at 04:08:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:08:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:08:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 cryogenic tank replacement +2025-04-04 at 04:08:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:08:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Failed oxygen tank mission commander Jim Lovell decision +2025-04-04 at 04:08:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-04 at 04:08:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:08:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 cryogenic tank failure prevention +2025-04-04 at 04:08:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:08:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 oxygen tank problem primary mission astronaut +2025-04-04 at 04:08:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-04 at 04:08:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:08:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew safety measures +2025-04-04 at 04:08:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:08:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:08:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew training and mission preparation +2025-04-04 at 04:08:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:08:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:08:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew resourcefulness in overcoming challenges +2025-04-04 at 04:08:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:08:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:08:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew adaptability and resourcefulness +2025-04-04 at 04:08:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:08:17 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:08:17 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:08:17 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, False, False, False] +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1984, 462, 740, 431, 312, 264] +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [84, 84, 84, 84, 84, 84] +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_correctness:84 - Average student length: 698.83 +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 84.00 +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_correctness:86 - Length ratio: 8.32 +2025-04-04 at 04:08:17 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-04 at 04:08:17 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.158 Âą 0.354 +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 0.83 Âą 1.86 +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 5, 0, 0, 0] +2025-04-04 at 04:08:17 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:08:17 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nDuring the flight, the pressure in the ascent stage oxygen tank 2 increased, ...', 'Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-04 at 04:08:17 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-04 at 04:08:17 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nBecause an inflight anomaly in the cryogenic oxygen supply required an abort ...'] +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 3.0 +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.500 +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.500 +2025-04-04 at 04:08:17 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.167, Perfect scores: 1/6 +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 1/6 +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 0.67 +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:08:17 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:08:17 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.248, Max reward: 0.601 +2025-04-04 at 04:08:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:08:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the reaction control system utilized in the lunar module of the Apollo missions? +2025-04-04 at 04:08:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:08:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system NASA Apollo 11 +2025-04-04 at 04:08:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control computer system chemistry control +2025-04-04 at 04:08:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Since it's related to an Apollo mission's lunar lander, let's see what it was established to do in lunar landings. +2025-04-04 at 04:08:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:08:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system +2025-04-04 at 04:08:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:08:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:08:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the primary function of the lunar module's reaction control system during descent and ascent phases? +2025-04-04 at 04:08:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:08:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module reaction control system +2025-04-04 at 04:08:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system chemistry control established +2025-04-04 at 04:08:24 | INFO | src.agent:run_tool_calls:118 - īŋŊīŋŊī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:08:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module Reaction Control System accomplished lunar orbit maneuvering and sampling functions. +2025-04-04 at 04:08:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system abort guidance +2025-04-04 at 04:08:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +With primary guidance available, guidance system steering was manually followed to trim the spacecraft attitudes for the maneuver. Although the displayed attitudes looked favorable in comparison to groundsupplied and out-the-window readings , the primary guidance steering needles read full scale left in roll and yaw (section 6.4). At about 137 hours 40 minutes, the lunar module reaction control system was used to provide a 2.9-ft/sec velocity correction. The maneuver was completed using manual pitch and roll control and abort guidance yaw control in a manner similar to that for the previous midcourse correction. + +8.9.4 Service Module Separation and Photography +------ + +2025-04-04 at 04:08:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:08:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the reaction control system's functions during the specific phases of lunar module operations: abort guidance and reaction control system check? +2025-04-04 at 04:08:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:08:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system NASA technical specifications +2025-04-04 at 04:08:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system attitude maneuvers +2025-04-04 at 04:08:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-04 at 04:08:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module Guidance Roll maneuver Follow thruster firing. +2025-04-04 at 04:08:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:08:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system primary function +2025-04-04 at 04:08:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:08:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the primary function of the reaction control system during the lunar module's entry into lunar orbit? +2025-04-04 at 04:08:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system undocking maneuvers +2025-04-04 at 04:08:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:08:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Capabilities of Lunar Module Reaction Control System during launch phase following lunar module separation. +2025-04-04 at 04:08:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module primary function Apollo +2025-04-04 at 04:08:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:08:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:08:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the primary function of the reaction control system during lunar module separation from the command module? +2025-04-04 at 04:08:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:08:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system attitude control modes +2025-04-04 at 04:08:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +attitude error needles to maintain attitude. Attitude control during the maneuver was performed by manually nulling the pitch and roll error needles. This maneuver necessarily required crew-cooperation, since the Lunar Module Pilot controlled pitch and the Commander controlled roll. Yaw attitude was maintained automatically by the abort guidance system. The Command Module Pilot called out the engine start and stop times, and the entire l4-second firing was performed at l0 percent thrust. The engine was shut down l second short of the calculated firing time to preclude an overburn which might require use of minus-X thrusters and cause plume impingement on the command module. The control and alignment techniques to accomplish such a contingency midcourse maneuver are believed to be satisfactory. +------ + +2025-04-04 at 04:08:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Operational success of Lunar Module Reaction Control System during separation and post-separation critical phases. +2025-04-04 at 04:08:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:08:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system documentation +2025-04-04 at 04:08:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:08:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system attitude attitude control +2025-04-04 at 04:08:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +attitude error needles to maintain attitude. Attitude control during the maneuver was performed by manually nulling the pitch and roll error needles. This maneuver necessarily required crew-cooperation, since the Lunar Module Pilot controlled pitch and the Commander controlled roll. Yaw attitude was maintained automatically by the abort guidance system. The Command Module Pilot called out the engine start and stop times, and the entire l4-second firing was performed at l0 percent thrust. The engine was shut down l second short of the calculated firing time to preclude an overburn which might require use of minus-X thrusters and cause plume impingement on the command module. The control and alignment techniques to accomplish such a contingency midcourse maneuver are believed to be satisfactory. +------ + +2025-04-04 at 04:08:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What can be achieved with Lunar Module Reaction Control System after service module jettison. +2025-04-04 at 04:08:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system primary function +2025-04-04 at 04:08:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:08:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system crew cooperation +2025-04-04 at 04:08:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:08:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Reviews of the Lunar Module Reaction Control System's performance. +2025-04-04 at 04:08:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module after service module separation +2025-04-04 at 04:08:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:08:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:08:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system evaluation +2025-04-04 at 04:08:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module Reaction Control System performance evaluation during lunar descent. +2025-04-04 at 04:08:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system separation procedures +2025-04-04 at 04:08:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:08:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:08:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo reaction control system evaluation procedures +2025-04-04 at 04:08:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Outcome of the Descent Propulsion System evaluation of the Lunar Module Reaction Control System. +2025-04-04 at 04:08:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system entry procedures +2025-04-04 at 04:08:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-04 at 04:08:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:08:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system evaluation criteria +2025-04-04 at 04:08:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-04 at 04:08:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Outcome of the comprehensive evaluation of the Lunar Module Reaction Control System. +2025-04-04 at 04:08:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system power systems +2025-04-04 at 04:08:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:08:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system propellant consumption +2025-04-04 at 04:08:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:08:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: .rc system success in crucial evaluation +2025-04-04 at 04:08:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module temporary electrical power system fault +2025-04-04 at 04:08:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-04 at 04:08:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:08:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system fuel usage analysis +2025-04-04 at 04:08:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-04 at 04:08:49 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:08:49 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:08:49 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, True, False] +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_correctness:82 - Student lengths: [690, 343, 1615, 863, 147, 2031] +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [28, 28, 28, 28, 28, 28] +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_correctness:84 - Average student length: 948.17 +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 28.00 +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_correctness:86 - Length ratio: 33.86 +2025-04-04 at 04:08:49 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:08:49 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.237 Âą 0.363 +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 6.83 Âą 13.17 +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [5, 0, 0, 36, 0, 0] +2025-04-04 at 04:08:49 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:08:49 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...'] +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:49 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...'] +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +attitude error needles to maintain attitude. Attitude control during the maneuver was performed by manually nulling the pitch and roll error needles. This maneuver necessarily required crew-cooperation, since the Lunar Module Pilot controlled pitch and the Commander controlled roll. Yaw attitude was maintained automatically by the abort guidance system. The Command Module Pilot called out the engine start and stop times, and the entire l4-second firing was performed at l0 percent thrust. The engine was shut down l second short of the calculated firing time to preclude an overburn which might require use of minus-X thrusters and cause plume impingement on the command module. The control and alignment techniques to accomplish such a contingency midcourse maneuver are believed to be satisfactory. +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +attitude error needles to maintain attitude. Attitude control during the maneuver was performed by manually nulling the pitch and roll error needles. This maneuver necessarily required crew-cooperation, since the Lunar Module Pilot controlled pitch and the Commander controlled roll. Yaw attitude was maintained automatically by the abort guidance system. The Command Module Pilot called out the engine start and stop times, and the entire l4-second firing was performed at l0 percent thrust. The engine was shut down l second short of the calculated firing time to preclude an overburn which might require use of minus-X thrusters and cause plume impingement on the command module. The control and alignment techniques to accomplish such a contingency midcourse maneuver are believed to be satisfactory. +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 12: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-04 at 04:08:49 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...', 'Result 1:\nThe reaction control system was activated at about 58 hours. Total propellant...', 'Result 1:\nService module.- At the time the system was powered down, reaction control sy...'] +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:49 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...'] +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-04 at 04:08:49 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +With primary guidance available, guidance system steering was manually followed to trim the spacecraft attitudes for the maneuver. Although the displayed attitudes looked favorable in comparison to groundsupplied and out-the-window readings , the primary guidance steering needles read full scale left in roll and yaw (section 6.4). At about 137 hours 40 minutes, the lunar module reaction control system was used to provide a 2.9-ft/sec velocity correction. The maneuver was completed using manual pitch and roll control and abort guidance yaw control in a manner similar to that for the previous midcourse correction. + +8.9.4 Service Module Separation and Photography +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-04 at 04:08:49 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nA.2 LUNAR MODULE\n------\nResult 2:\n1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 ...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nThe service module was separated $\\scriptstyle4-3/4$ hours before entry , aff...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...'] +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 0.0 +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.000 +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.000 +2025-04-04 at 04:08:49 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.333, Perfect scores: 0/6 +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 3.17 +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 0/6 +2025-04-04 at 04:08:49 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:08:49 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.208, Max reward: 0.561 +2025-04-04 at 04:08:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:08:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Southwest Airlines flight 842 diversion to Phoenix, Arizona +2025-04-04 at 04:08:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-04 at 04:08:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "crash Pago Pago disappearing flight" +2025-04-04 at 04:08:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-04 at 04:08:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what flight crew was flown from Pago Pago, Samoa to next destination +2025-04-04 at 04:08:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-04 at 04:08:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: American Airlines flight 33 rescue destination after Pago Pago +2025-04-04 at 04:08:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-04 at 04:08:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "flight crew Pago Pago Samoa destination +2025-04-04 at 04:08:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-04 at 04:08:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:08:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 intent to land South Pacific +2025-04-04 at 04:08:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:08:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "flight crew Iwo Jima Pago Pago" +2025-04-04 at 04:08:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-04 at 04:08:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: .where did the crew of a recent flight from Pago Pago, Samoa take a connecting flight to +2025-04-04 at 04:08:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +The first reported electronic contact by the recovery forces was through S-band contact by Samoa Rescue 4. A visual sighting report by the Recovery helicopter was received and was followed shortly thereafter by aquisition of the recovery beacon signal by the Recovery, Photo, and Swim l helicopters. Fuel dump was noted and voice contact was made with the descending spacecraft, although no latitude and longitude data were received. The command module landed at 1807 G.m.t. and remained in the stable l flotation attitude. The flashing light was operating and the infiation of the uprighting system commenced about l0 minutes subsequent to landing. +------ + +2025-04-04 at 04:08:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +2025-04-04 at 04:08:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:08:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 recovery operations flight crew destination +2025-04-04 at 04:08:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The Department of Defense provided recovery support commensurate with mission planning for Apollo l3. Because of the emergency which resulted in premature termination of the mission, additional support was provided by the Department of Defense and offers of assistance were made by many foreign nations, including England, France, Greece, Spain, Germany, Uruguay, Brazil, Kenya, the Netherlands, Nationalist China, and the Soviet Union. As a result of this voluntary support, a total of 2l ships and 17 aircraft were available for supporting an Indian Ocean landing, and 5l ships and 2l aircraft for an Atlantic Ocean landing. In the Pacific Ocean, there were l3 ships and 17 aircraft known to be available over and above the forces designated for primary recovery support. +------ + +2025-04-04 at 04:08:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:08:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew training and mission planning +2025-04-04 at 04:08:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:08:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "flight crew post recovery operations Iwo Jima" +2025-04-04 at 04:08:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-04 at 04:08:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the flight from Pago Pago, Samoa where the crew went next after returning from Hickam Air Force Base in Hawaii +2025-04-04 at 04:08:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-04 at 04:08:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "American Airlines flight 33 recovery operations Honolulu" +2025-04-04 at 04:08:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-04 at 04:08:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 10 command module recovery location +2025-04-04 at 04:08:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:08:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:09:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: low hydrogen pressure alarm 38 hours Apollo 13 +2025-04-04 at 04:09:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:09:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "icc NCO l00 hours" +2025-04-04 at 04:09:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew reported sleeping well the first 2 days of the mission. They all slept about 5-l/2 hours during the first sleep period. During the second period, the Commander, Command Module Pilot, and Lunar Module Pilot slept 5, 6, and 9 hours, respectively. The third sleep period was scheduled for 6l hours, but the oxygen tank incident at 56 hours precluded sleep by any of the crew until approximately 8o hours. +------ +Result 2: +The installed primary lithium hydroxide cartridge was used for approximately 27 hours (82 man-hoursīŧ‰ following activation of the lunar moduie at about 58 hours. The secondary cartridge was selected at about $\upbeta5$ 1/2 hours. During operation of the lunar module carbon dioxide removal system, the level was permitted to increase to an indicated l4.9 mm Hg. The primary cartridge is nominally rated for a usage capacity of 4l manhours at 520 Btu/man-hour. The secondary cartridge, nominally rated for about 17.9 man-hours, was used for 8 1/2 hours (25 1/2 man-hours). This cartridge is identical to that used in the portable life support system. A second primary cartridge was installed and used for approximately 6 minutes, but for the remainder of the mission, command module lithium hyaroxide cartridges were operated in a special arrangement. One side of each of two command module cartridges was covered and sealed with a plastic bag normally used to store a liquid-cooling garment. As shown +------ + +2025-04-04 at 04:09:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: post recovery inspection Iwo Jima to where in Hawaii +2025-04-04 at 04:09:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-04 at 04:09:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +2025-04-04 at 04:09:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 10 post-flight recovery operations +2025-04-04 at 04:09:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:09:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:09:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 recovery flights after Pago Pago" +2025-04-04 at 04:09:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:09:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mission briefing what was the flight plan for the Iwo Jima crew after departing Hawaii +2025-04-04 at 04:09:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-04 at 04:09:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +2025-04-04 at 04:09:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "when was the Apollo 10 command module recovered from the ocean, and where was it taken after recovery +2025-04-04 at 04:09:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:09:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:09:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 post-flight flights after Pago Pago" +2025-04-04 at 04:09:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:09:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: why did the Iwo Jima arrive in Hawaii in April 24 +2025-04-04 at 04:09:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-04 at 04:09:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +2025-04-04 at 04:09:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "USS Iwo Jima crew recovery operation +2025-04-04 at 04:09:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-04 at 04:09:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:09:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 crew flight after Pago Pago Samoa" +2025-04-04 at 04:09:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:09:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +2025-04-04 at 04:09:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:09:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 mission after return to Earth" +2025-04-04 at 04:09:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:09:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +2025-04-04 at 04:09:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:09:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 14 mission" +2025-04-04 at 04:09:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-04 at 04:09:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +2025-04-04 at 04:09:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:09:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 recovery crew after Pago Pago" +2025-04-04 at 04:09:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:09:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +2025-04-04 at 04:09:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:09:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +2025-04-04 at 04:09:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:09:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lat = 21.3182, long = -159.8123 +2025-04-04 at 04:09:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:09:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:09:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lat = 21.3182, long = -159.8123, date = March 29, 2023, time = 07:17 +2025-04-04 at 04:09:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:09:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:09:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +2025-04-04 at 04:09:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:09:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lat = 21.3182, long = -159.8123 +2025-04-04 at 04:09:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:09:19 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:09:19 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:09:19 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, False, False, False] +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_correctness:82 - Student lengths: [59, 1984, 407, 1159, 189, 663] +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [6, 6, 6, 6, 6, 6] +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_correctness:84 - Average student length: 743.50 +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 6.00 +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_correctness:86 - Length ratio: 123.92 +2025-04-04 at 04:09:19 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:09:19 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.396 Âą 0.426 +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 6.83 Âą 10.27 +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 6, 29, 6, 0] +2025-04-04 at 04:09:19 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:09:19 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe transearth injection maneuver was performed on time, and the transearth c...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The crew reported sleeping well the first 2 days of the mission. They all slept about 5-l/2 hours during the first sleep period. During the second period, the Commander, Command Module Pilot, and Lunar Module Pilot slept 5, 6, and 9 hours, respectively. The third sleep period was scheduled for 6l hours, but the oxygen tank incident at 56 hours precluded sleep by any of the crew until approximately 8o hours. +------ +Result 2: +The installed primary lithium hydroxide cartridge was used for approximately 27 hours (82 man-hoursīŧ‰ following activation of the lunar moduie at about 58 hours. The secondary cartridge was selected at about $\upbeta5$ 1/2 hours. During operation of the lunar module carbon dioxide removal system, the level was permitted to increase to an indicated l4.9 mm Hg. The primary cartridge is nominally rated for a usage capacity of 4l manhours at 520 Btu/man-hour. The secondary cartridge, nominally rated for about 17.9 man-hours, was used for 8 1/2 hours (25 1/2 man-hours). This cartridge is identical to that used in the portable life support system. A second primary cartridge was installed and used for approximately 6 minutes, but for the remainder of the mission, command module lithium hyaroxide cartridges were operated in a special arrangement. One side of each of two command module cartridges was covered and sealed with a plastic bag normally used to store a liquid-cooling garment. As shown +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +The first reported electronic contact by the recovery forces was through S-band contact by Samoa Rescue 4. A visual sighting report by the Recovery helicopter was received and was followed shortly thereafter by aquisition of the recovery beacon signal by the Recovery, Photo, and Swim l helicopters. Fuel dump was noted and voice contact was made with the descending spacecraft, although no latitude and longitude data were received. The command module landed at 1807 G.m.t. and remained in the stable l flotation attitude. The flashing light was operating and the infiation of the uprighting system commenced about l0 minutes subsequent to landing. +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 12: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 13: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 14: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 15: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The Department of Defense provided recovery support commensurate with mission planning for Apollo l3. Because of the emergency which resulted in premature termination of the mission, additional support was provided by the Department of Defense and offers of assistance were made by many foreign nations, including England, France, Greece, Spain, Germany, Uruguay, Brazil, Kenya, the Netherlands, Nationalist China, and the Soviet Union. As a result of this voluntary support, a total of 2l ships and 17 aircraft were available for supporting an Indian Ocean landing, and 5l ships and 2l aircraft for an Atlantic Ocean landing. In the Pacific Ocean, there were l3 ships and 17 aircraft known to be available over and above the forces designated for primary recovery support. +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-04 at 04:09:19 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:09:19 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.433, Perfect scores: 2/6 +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 2.17 +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:09:19 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:09:19 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.339, Max reward: 0.700 +2025-04-04 at 04:09:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:09:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space spacecraft brech plenum material +2025-04-04 at 04:09:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:09:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Advanced materials used in aerospace thermal insulation and their thermal conductivity in extreme temperature applications. +2025-04-04 at 04:09:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-04 at 04:09:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: advanced lightweight materials for rocket breech plenum chamber applications in future space missions +2025-04-04 at 04:09:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-04 at 04:09:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "spacecraft breech plenum material selection" +2025-04-04 at 04:09:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:09:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:09:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Insulation materials for components exposed to cryogenic temperatures in lunar missions +2025-04-04 at 04:09:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:09:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: propellant types used in spacecraft +2025-04-04 at 04:09:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-04 at 04:09:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: sheet material for spacecraft breech plenum +2025-04-04 at 04:09:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:09:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:09:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module structures and materials used for thermal insulation +2025-04-04 at 04:09:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:09:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: materials used in liquid hydrogen propellant systems +2025-04-04 at 04:09:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-04 at 04:09:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: polyimide vs other materials for spacecraft interior components +2025-04-04 at 04:09:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-04 at 04:09:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:09:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module heat shield or thermal insulation materials +2025-04-04 at 04:09:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-04 at 04:09:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: materials used in O-rings for cryogenic applications +2025-04-04 at 04:09:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-04 at 04:09:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spray foam insulation vs other materials for spacecraft components +2025-04-04 at 04:09:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:09:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:09:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Thermal insulation materials used in Apollo lunar modules +2025-04-04 at 04:09:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-04 at 04:09:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: materials with high-temperature and cryogenic resistance +2025-04-04 at 04:09:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:09:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aerospace materials used for spacecraft breech plenum +2025-04-04 at 04:09:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:09:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:09:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft thermal insulation materials and heat shield designs +2025-04-04 at 04:09:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:09:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: materials used for O-ring seals in cryogenic environments +2025-04-04 at 04:09:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-04 at 04:09:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:09:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft thermal insulation materials and specific component materials +2025-04-04 at 04:09:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-04 at 04:09:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:09:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft thermal insulation for component breech plenum area +2025-04-04 at 04:09:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:09:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:09:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Insulation materials used on lunar module components such as breech assemblies +2025-04-04 at 04:09:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-04 at 04:09:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:09:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Insulation materials used in lunar module components such as breech plenum areas +2025-04-04 at 04:09:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:09:45 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:09:45 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:09:45 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, False, True, False] +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_correctness:82 - Student lengths: [328, 1406, 1147, 2566, 244, 536] +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [9, 9, 9, 9, 9, 9] +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_correctness:84 - Average student length: 1037.83 +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 9.00 +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_correctness:86 - Length ratio: 115.31 +2025-04-04 at 04:09:45 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-04 at 04:09:45 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.237 Âą 0.363 +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.00 Âą 2.89 +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 7, 0, 5, 0] +2025-04-04 at 04:09:45 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:09:45 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...'] +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:09:45 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...', 'Result 1:\nThe Apollo lunar surface experiment package stowed for Apollo l3 was similar ...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...', 'Result 1:\nThe Apollo lunar surface experiment package stowed for Apollo l3 was similar ...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...'] +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-04 at 04:09:45 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe experiment consisted of an electronics package with sunshield and reflect...', 'Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...', 'Result 1:\nService module.- At the time the system was powered down, reaction control sy...', 'Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...', 'Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...', 'Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...'] +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-04 at 04:09:45 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-04 at 04:09:45 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 1.0 +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.167 +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:09:45 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.167, Perfect scores: 1/6 +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 1/6 +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 0.67 +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:09:45 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:09:45 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.370, Max reward: 0.837 +2025-04-04 at 04:09:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:09:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Verification task in Mars mission simulator for crew +2025-04-04 at 04:09:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:09:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: verified task space simulator crew preparation +2025-04-04 at 04:09:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:09:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: verifying tasks for astronauts before launch +2025-04-04 at 04:09:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:09:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA simulator task verification for crew +2025-04-04 at 04:09:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:09:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "verified task simulator space mission prep" +2025-04-04 at 04:09:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:09:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Simulation task Apollo crew preparation" +2025-04-04 at 04:09:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:09:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:09:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 mission pre-entry procedures verification +2025-04-04 at 04:09:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-04 at 04:09:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission task verification procedures for crew +2025-04-04 at 04:09:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:09:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA pre-entry procedures in simulator prior to advising crew +2025-04-04 at 04:09:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:09:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 mission crew simulator task" +2025-04-04 at 04:09:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:09:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: " Apollo mission simulation training procedures" +2025-04-04 at 04:09:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:09:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:09:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 lunar module entry, pre- and post-entry checks +2025-04-04 at 04:09:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:09:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 mission separation procedures +2025-04-04 at 04:09:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:09:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Pre-entry checks for lunar module in Apollo mission +2025-04-04 at 04:09:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:09:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 lunar module separation training" +2025-04-04 at 04:09:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:09:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 lunar landing simulation objective" +2025-04-04 at 04:09:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:09:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:10:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 mission task prior to Lunar Module re-entry +2025-04-04 at 04:10:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:10:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 entry procedures simulator validation +2025-04-04 at 04:10:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:10:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Pre-entry preparation for lunar module crew +2025-04-04 at 04:10:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-04 at 04:10:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 spacecraft re-entry procedure" +2025-04-04 at 04:10:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:10:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 12 mission objectives" +2025-04-04 at 04:10:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:10:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:10:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 mission procedure prior to Lunar Module re-entry +2025-04-04 at 04:10:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:10:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo inner spacecraft attitude control procedures +2025-04-04 at 04:10:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:10:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo command module power-up procedures for entry +2025-04-04 at 04:10:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:10:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 midcourse correction procedure" +2025-04-04 at 04:10:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:10:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 12 post-mission analysis" +2025-04-04 at 04:10:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:10:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:10:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 pre- and post-entry procedures +2025-04-04 at 04:10:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:10:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection and docking procedures +2025-04-04 at 04:10:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-04 at 04:10:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo simulator verification of pre-entry procedures +2025-04-04 at 04:10:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-04 at 04:10:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 midcourse correction maneuver" +2025-04-04 at 04:10:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:10:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 post-mission analysis" +2025-04-04 at 04:10:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:10:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:10:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 lunar module crew preparations for emergency return +2025-04-04 at 04:10:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:10:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar operations validation +2025-04-04 at 04:10:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-04 at 04:10:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module entry attitude and platform alignment +2025-04-04 at 04:10:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-04 at 04:10:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 spacecraft navigation" +2025-04-04 at 04:10:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:10:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 10 postflight analysis" +2025-04-04 at 04:10:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:10:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:10:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission requirements before launch +2025-04-04 at 04:10:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:10:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection +2025-04-04 at 04:10:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-04 at 04:10:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module platform alignment and misalignment +2025-04-04 at 04:10:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:10:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 navigation analysis" +2025-04-04 at 04:10:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:10:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 10 crew training and simulation" +2025-04-04 at 04:10:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:10:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:10:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 pre-launch training and helium pressure issues +2025-04-04 at 04:10:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:10:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo platform alignment tolerance +2025-04-04 at 04:10:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-04 at 04:10:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 post-flight evaluation" +2025-04-04 at 04:10:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:10:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:10:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 descent engine performance and hydrogen pressure management +2025-04-04 at 04:10:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:10:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo platform alignment standard +2025-04-04 at 04:10:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-04 at 04:10:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo mission report" +2025-04-04 at 04:10:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:10:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:10:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module systems simulation and procedures +2025-04-04 at 04:10:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:10:20 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:10:20 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:10:20 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1773, 605, 2026, 1150, 2033, 2038] +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_correctness:84 - Average student length: 1604.17 +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_correctness:86 - Length ratio: 160.42 +2025-04-04 at 04:10:20 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:10:20 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.237 Âą 0.237 +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 8.33 Âą 8.92 +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 3/6 +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 11, 17, 0, 22] +2025-04-04 at 04:10:20 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 6.0 +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 1.000 +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.000 +2025-04-04 at 04:10:20 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.400, Perfect scores: 2/6 +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.67 +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:10:20 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:10:20 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.386, Max reward: 0.724 +2025-04-04 at 04:10:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:10:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nachfol-end-to-end pressure measurement vacuum pumpdown +2025-04-04 at 04:10:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:10:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "vacuum pumpdown pressure measurement time frame" +2025-04-04 at 04:10:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-04 at 04:10:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long after pumpdown is jacket pressure measured to verify vacuum integrity in vacuum sealing applications +2025-04-04 at 04:10:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-04 at 04:10:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure measurement uniformity in a cryogenic pump-down process +2025-04-04 at 04:10:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-04 at 04:10:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long after pumpdown is pressure in the jacket measured to verify vacuum integrity in vacuum pumps +2025-04-04 at 04:10:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-04 at 04:10:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:10:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space shuttle helmet pressurization +2025-04-04 at 04:10:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-04 at 04:10:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "space suit pressure measurement timing" +2025-04-04 at 04:10:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:10:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "com NASA astronaut jacket pressure measurement time after pumpdown +2025-04-04 at 04:10:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:10:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: week-long vacuum integrity jacket pressure verification +2025-04-04 at 04:10:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-04 at 04:10:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long after pumpdown is direct jacket pressure measured to verify vacuum integrity +2025-04-04 at 04:10:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-04 at 04:10:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:10:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo space program extravehicular activity pressure garment assembly measurements +2025-04-04 at 04:10:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:10:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo command module and lunar module cabin pressure" +2025-04-04 at 04:10:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:10:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA astronaut jacket pressure measurement schedule post-launch" or "Astronaut suit pressure measurement timing after pumpdown in NASA missions +2025-04-04 at 04:10:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:10:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: jacket pressure measurement timing after pumpdown +2025-04-04 at 04:10:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:10:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the typical time after pumpdown to measure jacket pressure to ensure vacuum integrity +2025-04-04 at 04:10:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:10:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:10:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo space program pressure measurement methodology +2025-04-04 at 04:10:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:10:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "transient pressure behavior of supercritical fluid" +2025-04-04 at 04:10:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ +Result 2: +At about l09 hours when the helium bottle pressure had reached approximately l937 psi, the burst diaphragm ruptured and relieved the supercritical system through a special non-propulsive vent. The predicted rupture range for this vehicle was 190o Âą 20 psia. During venting, unexpected motion was inparted to the spacecraft which disrupted the motion established for the passive thermal control mode. The vent tube for the supercritical helium tank is ported on two sides by diametrically opposed Oval-shaped holes. It was originally believed that the escaping gas would exit these holes at 9o degrees to the tube axis such that no net thrust is produced. However, the pressure distribution in the tube is such that the two gas plumes have an included angle less than 180 degrees and probably closer to 9o degrees. Therefore, the component of the gas flow along the axis of the vent tube produces a net thrust in the opposite direction which tends to induce a slight roll rate to the vehicle. Since +------ + +2025-04-04 at 04:10:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: astronaut suit pressure measurement after launch timing onwards +2025-04-04 at 04:10:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:10:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure measurement in jacket after pumpdown verification +2025-04-04 at 04:10:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-04 at 04:10:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long after start of pumpdown is jacket pressure measured in vacuum pump systems +2025-04-04 at 04:10:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-04 at 04:10:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:10:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 helium tank pressure measurement timing +2025-04-04 at 04:10:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:10:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "aircraft instrumented thermal control" +2025-04-04 at 04:10:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:10:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA astronaut suit vacuum jacket pressure measurement timing documentation 1960s +2025-04-04 at 04:10:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-04 at 04:10:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: specific timing of jacket pressure measurement after pumpdown +2025-04-04 at 04:10:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:10:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Vacuum pump jacket pressure measurement time after pumpdown +2025-04-04 at 04:10:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:10:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:10:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ground-launched apollo space mission helium tank pressure measurement +2025-04-04 at 04:10:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:10:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Cabin and crew temperature control during launch and return phases" +2025-04-04 at 04:10:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-04 at 04:10:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA astronaut suit pressure measurement lag during lift-off +2025-04-04 at 04:10:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:10:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pumpdown jacket pressure measurement timing +2025-04-04 at 04:10:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-04 at 04:10:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Vacuum pump jacket pressure measurement time in pumps +2025-04-04 at 04:10:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-04 at 04:10:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:10:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module helium tank pressure measurement verification +2025-04-04 at 04:10:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:10:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Thermal stability in Command and Service Module (CSM)" +2025-04-04 at 04:10:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-04 at 04:10:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 2-3 week jacket pressure measurement timing +2025-04-04 at 04:10:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-04 at 04:10:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Vacuum system pressure measurement delay +2025-04-04 at 04:10:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-04 at 04:10:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:10:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank temperature range pressure measurement +2025-04-04 at 04:10:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:10:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Anomalies and issues during mission operation" +2025-04-04 at 04:10:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-04 at 04:10:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 2 or 3 week jacket pressure measurement +2025-04-04 at 04:10:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-04 at 04:10:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long after pumpdown is jacket pressure verified +2025-04-04 at 04:10:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:10:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:10:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank pressure measurement timing after pumpdown +2025-04-04 at 04:10:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:10:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Specific time frame for measuring jacket pressure after pumpdown in vacuum systems +2025-04-04 at 04:10:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-04 at 04:10:48 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:10:48 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:10:48 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, False, True, False] +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1847, 1797, 132, 229, 200, 1862] +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_correctness:84 - Average student length: 1011.17 +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 12.00 +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_correctness:86 - Length ratio: 84.26 +2025-04-04 at 04:10:48 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:10:48 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.317 Âą 0.354 +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 10.33 Âą 16.39 +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 2/6 +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 46, 6, 0, 10, 0] +2025-04-04 at 04:10:48 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ +Result 2: +At about l09 hours when the helium bottle pressure had reached approximately l937 psi, the burst diaphragm ruptured and relieved the supercritical system through a special non-propulsive vent. The predicted rupture range for this vehicle was 190o Âą 20 psia. During venting, unexpected motion was inparted to the spacecraft which disrupted the motion established for the passive thermal control mode. The vent tube for the supercritical helium tank is ported on two sides by diametrically opposed Oval-shaped holes. It was originally believed that the escaping gas would exit these holes at 9o degrees to the tube axis such that no net thrust is produced. However, the pressure distribution in the tube is such that the two gas plumes have an included angle less than 180 degrees and probably closer to 9o degrees. Therefore, the component of the gas flow along the axis of the vent tube produces a net thrust in the opposite direction which tends to induce a slight roll rate to the vehicle. Since +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-04 at 04:10:48 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 5.0 +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.833 +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:10:48 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.400, Perfect scores: 0/6 +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 4.17 +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 0/6 +2025-04-04 at 04:10:48 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:10:48 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.445, Max reward: 0.774 +2025-04-04 at 04:10:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:10:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: passive thermal control temperature range in satellite antenna +2025-04-04 at 04:10:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-04 at 04:10:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna temperature range passive thermal control mode +2025-04-04 at 04:10:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-04 at 04:10:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Temperature range antenna passive thermal control mode +2025-04-04 at 04:10:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-04 at 04:10:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature range of passive thermal control mode antenna temperature range +2025-04-04 at 04:10:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-04 at 04:10:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:10:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature range active and passive thermal control +2025-04-04 at 04:10:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-04 at 04:10:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-band antenna temperature range passive thermal control mode +2025-04-04 at 04:10:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-04 at 04:10:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature range antenna passive thermal control mode free return spacecraft +2025-04-04 at 04:10:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-04 at 04:10:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:10:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: passive thermal control mode temperature range vs engine firings +2025-04-04 at 04:10:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-04 at 04:10:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-wave antenna temperature requirements +2025-04-04 at 04:10:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:10:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna temperature range passive thermal control mode Âą$25^{\circ}$F +2025-04-04 at 04:10:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-04 at 04:10:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:10:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: passive thermal control mode, temperature range, 32:21:49 +2025-04-04 at 04:10:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-04 at 04:10:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-band antenna operation temperature range +2025-04-04 at 04:10:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ + +2025-04-04 at 04:10:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:11:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: free-returm midcourse correction 32:21:49 passive thermal control mode +2025-04-04 at 04:11:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-04 at 04:11:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:11:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna temperature range free-returm midcourse correction +2025-04-04 at 04:11:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-04 at 04:11:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:11:04 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:11:04 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:11:04 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, False, False, False, True] +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_correctness:82 - Student lengths: [333, 275, 303, 279, 466, 355] +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [29, 29, 29, 29, 29, 29] +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_correctness:84 - Average student length: 335.17 +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 29.00 +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_correctness:86 - Length ratio: 11.56 +2025-04-04 at 04:11:04 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-04 at 04:11:04 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.400 Âą 0.409 +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.17 Âą 2.34 +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [6, 4, 0, 0, 0, 3] +2025-04-04 at 04:11:04 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ + +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-04 at 04:11:04 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-04 at 04:11:04 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:11:04 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.467, Perfect scores: 2/6 +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.67 +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:11:04 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:11:04 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.329, Max reward: 0.682 +2025-04-04 at 04:11:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:11:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the first number in +2025-04-04 at 04:11:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:11:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first number in list +2025-04-04 at 04:11:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:11:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first number in a numerical sequence +2025-04-04 at 04:11:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:11:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is first number in an unordered list +2025-04-04 at 04:11:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:11:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the first number in a list definition +2025-04-04 at 04:11:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:11:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:11:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first number in sports score +2025-04-04 at 04:11:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:11:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: list of numbers sorted by position +2025-04-04 at 04:11:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +Time, hr:min Range station MILA Bandpass plots or tabs Bilevels Computers words O'graph records isruslh recrds :ipecial plots or tabs Speeial suo.tooal From To 00:00 00:02 00:04 00:07 00:14 01:31 01:33 01:48 02:25 02:34 02:43 02:49 04:44 08:35 12:49 13:18 16:44 17 :15 20 :37 24:53 27:01 37:33 40:55 44:39 00:12 00:14 04:44 00:18 00:28 01:33 01:45 01:59 02:34 02:45 02:50 13:59 08:35 12:49 16:44 17 :12 20 :37 25:00 BDA MSFN VAN CYI GDS MILA CYI CRO HAW HAW GDS MSFN MSFN MSFN HSK MSFN MAD 27:01 MSFN 37:42 GDS 40:55 MSFN 42:47 HSK 44:38 MS FN 52:37 MSFN 58:39 CDS 58:39 MS FN 101:58 GDS 123:12 GDS X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ + +2025-04-04 at 04:11:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: irst number in a dataset +2025-04-04 at 04:11:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:11:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:11:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first number in 2023 NBA season +2025-04-04 at 04:11:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:11:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: values 101 58 +2025-04-04 at 04:11:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +-1.2 40.4 +0.4 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:11:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: clinical trial data first number +2025-04-04 at 04:11:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:11:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:11:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nba scores 2023 season 1st game +2025-04-04 at 04:11:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +(section ll.3). +------ + +2025-04-04 at 04:11:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: heart rate 68 beats per minute +2025-04-04 at 04:11:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-04 at 04:11:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:11:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the first game score in nba history +2025-04-04 at 04:11:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +-1.2 40.4 +0.4 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:11:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:11:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how long did xavier jackson's team win or lose first game with him +2025-04-04 at 04:11:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ +Result 2: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ + +2025-04-04 at 04:11:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:11:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: wikipedia apollo 14 +2025-04-04 at 04:11:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:11:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:11:18 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:11:18 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:11:18 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, False, True, False, False] +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_correctness:82 - Student lengths: [303, 223, 28, 138, 158, 84] +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [6, 6, 6, 6, 6, 6] +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_correctness:84 - Average student length: 155.67 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 6.00 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_correctness:86 - Length ratio: 25.94 +2025-04-04 at 04:11:18 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.667, Valid formats: 4.0/6 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-04 at 04:11:18 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.458 Âą 0.371 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.50 Âą 2.50 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [7, 3, 0, 1, 4, 0] +2025-04-04 at 04:11:18 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: ++21.0 +3.0 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: ++21.0 +3.0 +------ +Result 2: +(section ll.3). +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +-1.2 40.4 +0.4 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ +Result 2: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +Time, hr:min Range station MILA Bandpass plots or tabs Bilevels Computers words O'graph records isruslh recrds :ipecial plots or tabs Speeial suo.tooal From To 00:00 00:02 00:04 00:07 00:14 01:31 01:33 01:48 02:25 02:34 02:43 02:49 04:44 08:35 12:49 13:18 16:44 17 :15 20 :37 24:53 27:01 37:33 40:55 44:39 00:12 00:14 04:44 00:18 00:28 01:33 01:45 01:59 02:34 02:45 02:50 13:59 08:35 12:49 16:44 17 :12 20 :37 25:00 BDA MSFN VAN CYI GDS MILA CYI CRO HAW HAW GDS MSFN MSFN MSFN HSK MSFN MAD 27:01 MSFN 37:42 GDS 40:55 MSFN 42:47 HSK 44:38 MS FN 52:37 MSFN 58:39 CDS 58:39 MS FN 101:58 GDS 123:12 GDS X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +-1.2 40.4 +0.4 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-04 at 04:11:18 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: ++21.0 +3.0 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 5.0 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.833 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:11:18 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.600, Perfect scores: 2/6 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 5/6 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.00 +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 4/6 +2025-04-04 at 04:11:18 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:11:18 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.397, Max reward: 0.678 +2025-04-04 at 04:11:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:11:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what percentage of stitch holes had cracks extending from them on a window shade? +2025-04-04 at 04:11:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What percentage of stitch holes from a window shade had cracks extending from them? +2025-04-04 at 04:11:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What percentage of stitch holes in a window shade fabric showed cracks or defects? +2025-04-04 at 04:11:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of cracked stitch holes in window shades +2025-04-04 at 04:11:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:11:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module window shade stitch hole crack percentage +2025-04-04 at 04:11:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of stitch holes with cracks originating from them on window shades +2025-04-04 at 04:11:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:11:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 window shade stitch hole ratio +2025-04-04 at 04:11:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of potential fatal flaws accommodated during window shade fabrication process +2025-04-04 at 04:11:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:11:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 window shade stitch hole percentage +2025-04-04 at 04:11:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of crack-prone stitch holes in window shade fabric +2025-04-04 at 04:11:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:11:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: compliance with stitches or reinforcement expansion to accommodate increased ductility for improved crack propagation +2025-04-04 at 04:11:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ + +2025-04-04 at 04:11:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:11:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of stitches with cracks from 80% compared to original 14.2.5 Cracked Window Shade +2025-04-04 at 04:11:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:11:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of stitching material flaws associated with mylar tape in lunar module shades +2025-04-04 at 04:11:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-04 at 04:11:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:11:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of cracks on Aclar shade stitching +2025-04-04 at 04:11:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:11:33 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:11:33 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:11:33 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, True, True, False, False] +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_correctness:82 - Student lengths: [401, 128, 335, 171, 140, 616] +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_correctness:84 - Average student length: 298.50 +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_correctness:86 - Length ratio: 29.85 +2025-04-04 at 04:11:33 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:11:33 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.350 Âą 0.394 +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.17 Âą 2.97 +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 4, 1, 8, 0, 0] +2025-04-04 at 04:11:33 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ + +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-04 at 04:11:33 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-04 at 04:11:33 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:11:33 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.167, Perfect scores: 1/6 +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 1/6 +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 0.50 +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:11:33 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:11:33 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.221, Max reward: 0.576 +2025-04-04 at 04:11:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:11:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the average fuel consumption of a liquid fuel service propulsion system like the one used on the International Space Station's Zvezda service module or the liquid-fueled rocket engines used in NASA's Apollo missions? +2025-04-04 at 04:11:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:11:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +spacecraft propulsion system fuel consumption Apollo + +2025-04-04 at 04:11:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:11:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space shuttle hubble space telescope fuel consumption +2025-04-04 at 04:11:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:11:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +.service propulsion system fuel consumption +2025-04-04 at 04:11:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-04 at 04:11:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Fuel consumption of a service propulsion system for a lunar sample return mission +2025-04-04 at 04:11:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:11:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:11:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the fuel consumption of the Service Propulsion System (SPS) for the International Space Station? +2025-04-04 at 04:11:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:11:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Apollo 11 Descent Propulsion System propellant consumption + +2025-04-04 at 04:11:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:11:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo mission fuel consumption +2025-04-04 at 04:11:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:11:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +apollo 15 propellant consumptions +2025-04-04 at 04:11:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:11:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Descent Propulsion System Fuel Consumption +2025-04-04 at 04:11:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:11:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:11:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the total propellant consumption of the Space Shuttle's main propulsion system during a mission? +2025-04-04 at 04:11:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-04 at 04:11:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 systems power consumption +2025-04-04 at 04:11:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:11:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +apollo 17 propellant consumptions +2025-04-04 at 04:11:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-04 at 04:11:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Descent Propulsion System Propellant Mass Consumption without actually consuming propellants +2025-04-04 at 04:11:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-04 at 04:11:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:11:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the total fuel consumption of the Space Shuttle's Service Propulsion System during a mission? +2025-04-04 at 04:11:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:11:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar module power consumption +2025-04-04 at 04:11:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:11:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +apollo 17 service vs descent propulsion fuel consumptions +2025-04-04 at 04:11:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:11:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Descent Propulsion System Total Mass Consumption +2025-04-04 at 04:11:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:11:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:11:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 descent phase fuel consumption +2025-04-04 at 04:11:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:11:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Total Consumed Descent Propulsion System Propellants Apollo 11 +2025-04-04 at 04:11:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:11:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:11:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar reaction control heaters power +2025-04-04 at 04:11:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:11:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:11:56 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:11:56 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:11:56 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, False, False] +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_correctness:82 - Student lengths: [431, 421, 1234, 489, 1088, 500] +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_correctness:84 - Average student length: 693.83 +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 4.00 +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_correctness:86 - Length ratio: 173.46 +2025-04-04 at 04:11:56 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-04 at 04:11:56 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.346 Âą 0.362 +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.33 Âą 2.43 +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [4, 0, 0, 0, 4, 6] +2025-04-04 at 04:11:56 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:11:56 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nTotal oxygen usage from the three lunar module oxygen tanks was 20.3 pounds o...', 'Result 1:\nService module.- At the time the system was powered down, reaction control sy...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...', 'Result 1:\nService module.- At the time the system was powered down, reaction control sy...'] +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:11:56 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...'] +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:11:56 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nTotal oxygen usage from the three lunar module oxygen tanks was 20.3 pounds o...', 'Result 1:\nThe crew rationed water and used it sparingly after the oxygen tank incident....', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-04 at 04:11:56 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:11:56 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...', 'Result 1:\nFuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5...', 'Result 1:\nSpacecraft mass properties for the Apollo l2 mission are summarized in table ...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...'] +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 1.0 +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.167 +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:11:56 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.167, Perfect scores: 0/6 +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 2/6 +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 0.83 +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 0/6 +2025-04-04 at 04:11:56 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:11:56 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.428, Max reward: 0.578 +2025-04-04 at 04:11:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:12:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: unmanned aerial vehicle fault +2025-04-04 at 04:12:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:12:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "SpaceX Starship unplanned minimum impulse" +2025-04-04 at 04:12:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-04 at 04:12:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "MINIATURE IMpulse SOLID-State Battery emergency firing 32:21:49" +2025-04-04 at 04:12:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:12:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: flight english tsa minpmf engine firing december 22 2022 +2025-04-04 at 04:12:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:12:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What triggered the unplanned minimum impulse engine firing at 32:21:49, specifically in the context of a jet engine? +2025-04-04 at 04:12:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-04 at 04:12:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:12:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 unplanned minimum impulse firing +2025-04-04 at 04:12:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:12:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "MSC-02680 Starship unplanned minimum impulse engine reason" +2025-04-04 at 04:12:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:12:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 16 lunar module descent engine minimum impulse firing thumping noise snowflakes" +2025-04-04 at 04:12:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-04 at 04:12:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the purpose and function of the roll manual attitude switch, specifically the difference between rate-command and acceleration-command positions? +2025-04-04 at 04:12:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-04 at 04:12:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:12:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 unplanned engine firing investigation +2025-04-04 at 04:12:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:12:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Starship 32:21:49 engine firing coning incorrect roll manual attitude switch" +2025-04-04 at 04:12:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-04 at 04:12:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 16 lunar module dc electrical system short in quadrant 4 venting" +2025-04-04 at 04:12:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:12:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the recommended operational procedure for switching from rate-command to acceleration-command position on the roll manual attitude switch? +2025-04-04 at 04:12:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ + +2025-04-04 at 04:12:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:12:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 unplanned engine firing +2025-04-04 at 04:12:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:12:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Starship September 2023 MSC-02680 ignition reason" +2025-04-04 at 04:12:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:12:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 16 lunar module descent battery short circuit components" +2025-04-04 at 04:12:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:12:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the purpose and design of the flight director attitude indicators in the context of Apollo missions? +2025-04-04 at 04:12:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:12:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:12:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 engine firing Apollo 13 +2025-04-04 at 04:12:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:12:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Starship 2023 MSC-02680 engine firing at 32:21:49" +2025-04-04 at 04:12:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:12:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 16 lunar module electrical power system l2-ampere mode" +2025-04-04 at 04:12:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:12:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the symptoms of an Unplanned Minimum Impulse Engine Firing and how might they affect the spacecraft's attitude control? +2025-04-04 at 04:12:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ + +2025-04-04 at 04:12:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:12:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: can someone describe the unplanned engine firing that caused Apollo 10's aborted lunar landing +2025-04-04 at 04:12:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-04 at 04:12:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Starship 2023 engine shutdown at 32:21:49" +2025-04-04 at 04:12:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-04 at 04:12:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 16 power mode comparison with Apollo 13" +2025-04-04 at 04:12:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:12:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the quad C isolation valves, and what was their effect on attitude control during the Unplanned Minimum Impulse Engine Firing? +2025-04-04 at 04:12:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:12:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:12:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 unplanned engine firing +2025-04-04 at 04:12:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:12:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "unplanned minimum impulse engine firing 2023 Starship post-Abort-12 mission analogy" +2025-04-04 at 04:12:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +4.0 TRAJECTORY + +The planned trajectory profile was similar to that for Apollo 12 except for descent orbit insertion being performed with the service propulsion system and the targeting of the spent S--IvB stage for a lunar impact. The trajectory had been very close to the nominal flight plan up to the time of abort, which was the first in the Apollo program. Throughout the manned space program, techniques have been developed and tested for the real-tine determination of immediate abort requirements, but Apollo l3 presented the first situation in which their use was necesSary. Figure 3-l shows the mission profile, including the relative locations of all major maneuvers. +------ + +2025-04-04 at 04:12:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 12 lunar module helium pressure increase" +2025-04-04 at 04:12:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:12:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the specific wiring issue discovered in the reaction control system, and how did it affect the operation of the propellant isolation valves? +2025-04-04 at 04:12:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:12:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:12:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 unplanned engine firing +2025-04-04 at 04:12:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:12:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Starship 2023 abort and S-II center engine shutdown" +2025-04-04 at 04:12:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-04 at 04:12:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Abnormal helium pressure rise rates in lunar module descent tanks" +2025-04-04 at 04:12:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:12:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:12:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Starship crew announced reason for minimum impulse engine firing" +2025-04-04 at 04:12:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:12:22 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:12:22 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:12:22 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, False, True] +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_correctness:82 - Student lengths: [516, 709, 1368, 1717, 500, 578] +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [112, 112, 112, 112, 112, 112] +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_correctness:84 - Average student length: 898.00 +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 112.00 +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_correctness:86 - Length ratio: 8.02 +2025-04-04 at 04:12:22 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-04 at 04:12:22 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.317 Âą 0.448 +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.50 Âą 3.55 +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [8, 0, 0, 0, 0, 7] +2025-04-04 at 04:12:22 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:12:22 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe extensive testing and analyses and the consistency with which the postfli...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nIgnition and lift-off occurred on schedule. A listing_of major flight plan ev...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nBecause an inflight anomaly in the cryogenic oxygen supply required an abort ...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...'] +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-04 at 04:12:22 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +4.0 TRAJECTORY + +The planned trajectory profile was similar to that for Apollo 12 except for descent orbit insertion being performed with the service propulsion system and the targeting of the spent S--IvB stage for a lunar impact. The trajectory had been very close to the nominal flight plan up to the time of abort, which was the first in the Apollo program. Throughout the manned space program, techniques have been developed and tested for the real-tine determination of immediate abort requirements, but Apollo l3 presented the first situation in which their use was necesSary. Figure 3-l shows the mission profile, including the relative locations of all major maneuvers. +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:12:22 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe data show that descent battery 2 experienced at least a 60-ampere current...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nAt 97 hours 14 minutes, the crew reported a thumping noise and snowflakes ven...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...'] +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:12:22 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nManeuver System Ignition time, hr:min:sec Firing time; sec Velocity change. f...'] +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 2.0 +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.333 +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:12:22 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.300, Perfect scores: 1/6 +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.33 +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:12:22 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:12:22 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.154, Max reward: 0.722 +2025-04-04 at 04:12:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:12:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 17 lunar module impact site coordinates, Apollo 17 landing site, seismic stations, seismic signal velocity +2025-04-04 at 04:12:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:12:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the distance from the S-IVB impact launch site to a seismic station? +2025-04-04 at 04:12:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:12:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 S-IVB impact velocity of initial signal to seismic station +2025-04-04 at 04:12:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:12:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the velocity of the seismic signal from S-IVB impact at Deep Impact station +2025-04-04 at 04:12:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:12:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: New Horizons Pluto flyby date +2025-04-04 at 04:12:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ + +2025-04-04 at 04:12:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:12:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 17 S-IVB impact signal velocity, moon structure, seismic signal characteristics +2025-04-04 at 04:12:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:12:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the fastest speed of a signal from a lunar impact to a seismic station? +2025-04-04 at 04:12:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:12:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: seismic velocity of S-IVB impact signal +2025-04-04 at 04:12:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:12:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Velocities of seismic signal from lunar impact +2025-04-04 at 04:12:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:12:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB impact velocity +2025-04-04 at 04:12:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:12:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:12:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 17 initial seismic signal velocity, Apollo 13 mission, seismic signal characteristics +2025-04-04 at 04:12:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:12:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: velocity of seismic signal from SIVB impact compared to ascent stage impact +2025-04-04 at 04:12:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-04 at 04:12:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB surface velocity at cutoff time +2025-04-04 at 04:12:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:12:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:12:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Escape velocity of lunar surface material, initial signal velocity, Apollo 17 +2025-04-04 at 04:12:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:12:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: peak intensity of SIVB impact signal compared to Apollo 12 ascent stage signal +2025-04-04 at 04:12:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:12:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:12:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 17 initial lunar module velocity, lunar orbital velocity, Apollo 17 mission +2025-04-04 at 04:12:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:12:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effects of S-IVB impact point on seismic signal +2025-04-04 at 04:12:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:12:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:12:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 17 guidance system, Apollo 17 navigation system, Apollo 17 initial signal velocity +2025-04-04 at 04:12:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:12:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: estimate distance to seismic station from S-IVB impact point +2025-04-04 at 04:12:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:12:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:12:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 17 initial signal velocity +2025-04-04 at 04:12:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:12:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: deviation of S-IVB impact point from planned trajectory +2025-04-04 at 04:12:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:12:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:12:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 17 initial signal velocity +2025-04-04 at 04:12:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:12:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: distance between S-IVB impact point and Apollo 12 seismometer +2025-04-04 at 04:12:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:12:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:12:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB impact point to Apollo 12 seismometer distance and entry time +2025-04-04 at 04:12:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:12:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:12:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB entry time and velocity at impact +2025-04-04 at 04:12:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:12:46 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:12:46 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:12:46 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, True, False, False, False] +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1766, 263, 121, 1206, 250, 175] +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_correctness:84 - Average student length: 630.17 +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_correctness:86 - Length ratio: 63.02 +2025-04-04 at 04:12:46 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.667, Valid formats: 4.0/6 +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:12:46 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.275 Âą 0.280 +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 1.17 Âą 1.21 +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 2, 2, 0, 0, 3] +2025-04-04 at 04:12:46 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-04 at 04:12:46 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:12:46 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe accuracy of the first midcourse correction (table 4-Il), which placed the...', 'Result 1:\nThe discarded S-IVB stage was targeted for a lunar impact of 3 degrees south ...', 'Result 1:\ncrewmen noted the small change in acceleration caused by the mixture ratio sh...'] +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:12:46 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.500, Perfect scores: 1/6 +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 2.00 +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:12:46 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:12:46 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.551, Max reward: 0.755 +2025-04-04 at 04:12:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:12:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: -orbital insertion delay space mission +2025-04-04 at 04:12:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:12:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: orbital insertion event time difference +2025-04-04 at 04:12:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ + +2025-04-04 at 04:12:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Space mission orbital insertion timing" +2025-04-04 at 04:12:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ + +2025-04-04 at 04:12:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the timing difference of an orbital insertion for the Dragon Capsule 2 mission to the orbital insertion point? +2025-04-04 at 04:12:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-04 at 04:12:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:12:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 orbital insertion delay +2025-04-04 at 04:12:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:12:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 entry time +2025-04-04 at 04:12:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:12:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Orbital insertion timing Apollo mission" +2025-04-04 at 04:12:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:12:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: when did the orbital insertion occur 44 seconds later than predicted +2025-04-04 at 04:12:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:12:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:12:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 orbital insertion delay time +2025-04-04 at 04:12:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:12:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 third midcourse correction +2025-04-04 at 04:12:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:12:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar module separation from command module" +2025-04-04 at 04:12:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-04 at 04:12:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: when was orbital insertion of the Dragon Capsule 2 mission +2025-04-04 at 04:12:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-04 at 04:12:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:12:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 orbital insertion time +2025-04-04 at 04:12:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:12:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 third midcourse correction time +2025-04-04 at 04:12:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:12:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar module separation time Apollo 12" +2025-04-04 at 04:12:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-04 at 04:12:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: when did the lunar module enter lunar orbit or complete orbital insertion +2025-04-04 at 04:12:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:12:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:13:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 mission duration +2025-04-04 at 04:13:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:13:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 predicted insertions and actual insertion +2025-04-04 at 04:13:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:13:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Orbital insertion timing versus lunar module separation time Apollo 12" +2025-04-04 at 04:13:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:13:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: when did the spacecraft enter lunar orbit +2025-04-04 at 04:13:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:13:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:13:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 orbital insertion time +2025-04-04 at 04:13:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:13:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 launch and third mission timeline +2025-04-04 at 04:13:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:13:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Spacecraft orbital insertion timing before lunar module separation Apollo 12" +2025-04-04 at 04:13:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:13:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: orbital insertion time for Apollo 11 was 142:30:00 +2025-04-04 at 04:13:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:13:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:13:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 range zero time to entry +2025-04-04 at 04:13:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:13:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar module power provision to command module before entry timing Apollo 12" +2025-04-04 at 04:13:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:13:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Transearth midcourse correction time for Apollo 11 +2025-04-04 at 04:13:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:13:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:13:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 range zero time entry time +2025-04-04 at 04:13:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:13:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar module power disconnection timing from command module Apollo 12" +2025-04-04 at 04:13:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:13:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:13:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 range zero to entry time +2025-04-04 at 04:13:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:13:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar module undocking from Apollo 12 at 11:11 minutes before entry" +2025-04-04 at 04:13:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:13:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:13:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar module separation timing Apollo 12" +2025-04-04 at 04:13:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-04 at 04:13:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:13:11 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:13:11 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:13:11 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_correctness:82 - Student lengths: [307, 2015, 194, 557, 631, 284] +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_correctness:84 - Average student length: 664.67 +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 5.00 +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_correctness:86 - Length ratio: 132.93 +2025-04-04 at 04:13:11 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-04 at 04:13:11 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.000 Âą 0.000 +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 0.00 Âą 0.00 +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 0, 0, 0, 0] +2025-04-04 at 04:13:11 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:13:11 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe planned launch and earth parking orbit phases for this mission were very ...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nIn this report, all actual times prior to earth landing are elapsed time from...', 'Result 1:\nIn this report, all actual times prior to earth landing are elapsed time from...', 'Result 1:\nIn this report, all actual times prior to earth landing are elapsed time from...'] +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-04 at 04:13:11 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe planned launch and earth parking orbit phases for this mission were very ...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nAt 2-l/2 hours prior to entry, the command module was fully powered up and lu...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nThe unprecedented powered-down state of the command module required generatio...', 'Result 1:\nAt 2-l/2 hours prior to entry, the command module was fully powered up and lu...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...'] +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-04 at 04:13:11 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-04 at 04:13:11 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:13:11 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe planned launch and earth parking orbit phases for this mission were very ...', 'Result 1:\nThe entry attitude and platform alignment were confirmed by a successful sext...', 'Result 1:\nAt approximately 105 hours, the crew performed a manual descent propulsion ma...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...'] +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 1.0 +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.167 +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:13:11 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.300, Perfect scores: 0/6 +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 2.67 +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 0/6 +2025-04-04 at 04:13:11 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:13:11 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.139, Max reward: 0.651 +2025-04-04 at 04:13:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:13:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar altitude unit +2025-04-04 at 04:13:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-04 at 04:13:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'lunar surface measurement unit' +2025-04-04 at 04:13:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:13:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: altitude unit lunar surface +2025-04-04 at 04:13:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-04 at 04:13:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar altitude measurement unit +2025-04-04 at 04:13:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-04 at 04:13:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: altitude unit lunar surface +2025-04-04 at 04:13:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-04 at 04:13:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar altitude unit" +2025-04-04 at 04:13:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:13:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:13:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar altitude measurement units +2025-04-04 at 04:13:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ + +2025-04-04 at 04:13:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'lunar altitude measurement unit' +2025-04-04 at 04:13:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ + +2025-04-04 at 04:13:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: altitude lunar surface definition +2025-04-04 at 04:13:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-04 at 04:13:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moon surface altitude measurement unit +2025-04-04 at 04:13:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-04 at 04:13:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mean lunar radius in feet +2025-04-04 at 04:13:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +Positions are extrapolated. + +The ll photographs have been normalized so that the earth is the same size in all frames. Frames 8590 and 8591 have been further enlarged. By viewing these two frames under a stereoscope, pronounced apparent relief is seen in the cloud patterns. The relief is so pronounced, in fact, that it cannot be attributed solely to height differences of clouds. It appears to result, in part, from the relative horizontal motion in the cloud fields; that is, clouds moving in the same direction as the spacecraft appear farther away than those moving in the direction opposite that of the spacecraftīŧŒ + +11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT +------ + +2025-04-04 at 04:13:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo lunar module navigation unit altitude measurement" +2025-04-04 at 04:13:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:13:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:13:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'lunar mean altitude unit' +2025-04-04 at 04:13:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-04 at 04:13:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar surface altitude above lunar radius +2025-04-04 at 04:13:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:13:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar surface experiment package altitude measurement unit +2025-04-04 at 04:13:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-04 at 04:13:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mean lunar radius in feet +2025-04-04 at 04:13:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +Positions are extrapolated. + +The ll photographs have been normalized so that the earth is the same size in all frames. Frames 8590 and 8591 have been further enlarged. By viewing these two frames under a stereoscope, pronounced apparent relief is seen in the cloud patterns. The relief is so pronounced, in fact, that it cannot be attributed solely to height differences of clouds. It appears to result, in part, from the relative horizontal motion in the cloud fields; that is, clouds moving in the same direction as the spacecraft appear farther away than those moving in the direction opposite that of the spacecraftīŧŒ + +11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT +------ + +2025-04-04 at 04:13:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "cruise altitude unit" +2025-04-04 at 04:13:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The structural evaluation is based on guidance and control data, cabin pressure measurements, conmand module acceleration data, photographs īŧŒ and crew comments . + +Based on measured command module accelerations and on simulations using actual launch wind data, lumar module loads were within structural limits during laumch and translurar injection. Loads during docking and service propulsion and descent prcpulsion maneuvers were also within structural limits. + +Data telemetered during the oxygen tank incident indicate the presence of body bending oscillations in the docked spacecraft. The associated amplitudes, however, were of a very low level, and bending loads in the critical docking-tmnel area were well below design limits. + +6.2 ELECTRICAL POWER +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:13:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:13:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'lunar mean radius altitude unit' +2025-04-04 at 04:13:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:13:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: altitude lunar mean radius +2025-04-04 at 04:13:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-04 at 04:13:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: heat flow experiment unit of measurement +2025-04-04 at 04:13:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-04 at 04:13:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module altitude measurement unit" +2025-04-04 at 04:13:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-04 at 04:13:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:13:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'lunar orbit altitude unit' +2025-04-04 at 04:13:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:13:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar mean radius in feet +2025-04-04 at 04:13:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:13:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollodata thermal gradient sensor unit +2025-04-04 at 04:13:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-04 at 04:13:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module inertial measurement unit" +2025-04-04 at 04:13:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:13:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:13:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mean lunar radius value +2025-04-04 at 04:13:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-04 at 04:13:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: thermocouple units +2025-04-04 at 04:13:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-04 at 04:13:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module inertial measurement unit altitude measurement" +2025-04-04 at 04:13:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-04 at 04:13:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:13:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar mean radius values +2025-04-04 at 04:13:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-04 at 04:13:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:13:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar radius in feet +2025-04-04 at 04:13:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-04 at 04:13:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:13:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module altitude reference +2025-04-04 at 04:13:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-04 at 04:13:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:13:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module altitude units +2025-04-04 at 04:13:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:13:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:13:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar surface altitude unit apollo +2025-04-04 at 04:13:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:13:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:13:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft altitude unit +2025-04-04 at 04:13:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:13:34 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:13:34 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:13:34 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, False, False, False, True] +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_correctness:82 - Student lengths: [386, 232, 2023, 257, 256, 249] +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_correctness:84 - Average student length: 567.17 +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 13.00 +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_correctness:86 - Length ratio: 43.63 +2025-04-04 at 04:13:34 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:13:34 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.396 Âą 0.426 +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 3.00 Âą 3.00 +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 6, 0, 6, 0, 6] +2025-04-04 at 04:13:34 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 12: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +Positions are extrapolated. + +The ll photographs have been normalized so that the earth is the same size in all frames. Frames 8590 and 8591 have been further enlarged. By viewing these two frames under a stereoscope, pronounced apparent relief is seen in the cloud patterns. The relief is so pronounced, in fact, that it cannot be attributed solely to height differences of clouds. It appears to result, in part, from the relative horizontal motion in the cloud fields; that is, clouds moving in the same direction as the spacecraft appear farther away than those moving in the direction opposite that of the spacecraftīŧŒ + +11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +Positions are extrapolated. + +The ll photographs have been normalized so that the earth is the same size in all frames. Frames 8590 and 8591 have been further enlarged. By viewing these two frames under a stereoscope, pronounced apparent relief is seen in the cloud patterns. The relief is so pronounced, in fact, that it cannot be attributed solely to height differences of clouds. It appears to result, in part, from the relative horizontal motion in the cloud fields; that is, clouds moving in the same direction as the spacecraft appear farther away than those moving in the direction opposite that of the spacecraftīŧŒ + +11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The structural evaluation is based on guidance and control data, cabin pressure measurements, conmand module acceleration data, photographs īŧŒ and crew comments . + +Based on measured command module accelerations and on simulations using actual launch wind data, lumar module loads were within structural limits during laumch and translurar injection. Loads during docking and service propulsion and descent prcpulsion maneuvers were also within structural limits. + +Data telemetered during the oxygen tank incident indicate the presence of body bending oscillations in the docked spacecraft. The associated amplitudes, however, were of a very low level, and bending loads in the critical docking-tmnel area were well below design limits. + +6.2 ELECTRICAL POWER +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 6.0 +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 1.000 +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.000 +2025-04-04 at 04:13:34 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.400, Perfect scores: 1/6 +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 2.17 +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:13:34 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:13:34 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.353, Max reward: 0.518 +2025-04-04 at 04:13:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:13:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the temperature of an antenna when its heaters are turned off in general antenna operation? +2025-04-04 at 04:13:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:13:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the typical temperature of an antenna when the antenna heaters are turned off for most common types of antennas used in electronic devices? +2025-04-04 at 04:13:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:13:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the operating temperature of the Mars Curiosity Rover's antenna when it was turned off during its Mars landing on August 6, 2012? +2025-04-04 at 04:13:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-04 at 04:13:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna heater shutdown temperature conditions +2025-04-04 at 04:13:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:13:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:13:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Temperature of antenna heaters turned off for a space-based antenna +2025-04-04 at 04:13:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-04 at 04:13:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the cabin temperature for a spacecraft when the onboard heaters are turned off? +2025-04-04 at 04:13:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:13:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna temperature after turning off antenna heaters on a spacecraft +2025-04-04 at 04:13:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-04 at 04:13:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:13:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Temperature of S-band steerable antenna heaters turned off in an operational spacecraft during space mission +2025-04-04 at 04:13:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:13:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the cabin temperature during landing after power has been off for an unspecified period? +2025-04-04 at 04:13:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:13:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna temperature passive thermal control +2025-04-04 at 04:13:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-04 at 04:13:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:13:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-band antenna thermal behavior during low-power communication mode in space mission +2025-04-04 at 04:13:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-04 at 04:13:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna heater shutdown temperature in space +2025-04-04 at 04:13:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-04 at 04:13:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:13:50 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:13:50 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:13:50 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, False, False, True, True] +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.67 +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_correctness:82 - Student lengths: [384, 563, 299, 898, 481, 859] +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [11, 11, 11, 11, 11, 11] +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_correctness:84 - Average student length: 580.67 +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 11.00 +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_correctness:86 - Length ratio: 52.79 +2025-04-04 at 04:13:50 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-04 at 04:13:50 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.375 Âą 0.378 +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 1.83 Âą 1.86 +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [4, 3, 0, 0, 4, 0] +2025-04-04 at 04:13:50 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-04 at 04:13:50 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-04 at 04:13:50 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:13:50 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 0/6 +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 0.00 +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 0/6 +2025-04-04 at 04:13:50 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:13:50 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.355, Max reward: 0.851 +2025-04-04 at 04:13:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:13:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: rise rate from lift-off to first descent propulsion maneuver Mars 202verance +2025-04-04 at 04:13:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:13:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: "Mars 2020 rover takeoff to landing descent propulsion average rise rate" +2025-04-04 at 04:13:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:13:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "average rise rate to first descent propulsion maneuver lunar mission" +2025-04-04 at 04:13:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:13:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: + Apollo 11 ascent rate +2025-04-04 at 04:13:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:13:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is Apollo 11 ascent and descent propulsion rates +2025-04-04 at 04:13:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:13:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the average rise rate from liftoff to first descent propulsion maneuver for Mars Curiosity Rover's landing process? +2025-04-04 at 04:13:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:13:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:13:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: rise rate Mars 2020 rover Perseverance +2025-04-04 at 04:13:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +boost phases were within acceptable spacecraft structural. design limits. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:13:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 mission lift-off time and average rise rate" +2025-04-04 at 04:13:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-04 at 04:13:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Apollo 11 ascent profile average rate of ascent during transfer from the Earth's surface to lunar orbit +2025-04-04 at 04:13:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:13:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 ascent and descent trajectory rates +2025-04-04 at 04:13:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:13:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (confidence: The information provided does not seem to match the Apollo 8 mission; it discusses a different lunar mission and doesn't explicitly state the spacecraft's system.) Apollo 8 was related to the moon, so I will try NASA's Apollo mission database. +2025-04-04 at 04:13:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-04 at 04:13:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:14:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: rise rate ascent trajectory Mars Perseverance rover +2025-04-04 at 04:14:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:14:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 lift-off performance parameters" +2025-04-04 at 04:14:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: + Apollo 11 ascent velocity at T+00:00 +2025-04-04 at 04:14:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:14:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 ascent and descent trajectory velocity change +2025-04-04 at 04:14:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:14:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the average rise rate from liftoff to surface descent for the first lunar module of the Apollo 8 mission? +2025-04-04 at 04:14:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:14:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:14:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Mars 2020 rover ascent trajectory +2025-04-04 at 04:14:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:14:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 descent propulsion system start time and rise rate" +2025-04-04 at 04:14:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:14:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 ascent and descent velocity change from range zero +2025-04-04 at 04:14:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:14:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Was the average rise rate from liftoff to first engine fire for the Service Module of Apollo 12? +2025-04-04 at 04:14:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-04 at 04:14:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:14:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mars 2020 ascent timeline +2025-04-04 at 04:14:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:14:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 actual first landing descent propulsion system start time" +2025-04-04 at 04:14:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:14:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:14:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Mars 2020 ascent trajectory details +2025-04-04 at 04:14:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:14:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 lift-off to first descent propulsion system start time" +2025-04-04 at 04:14:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:14:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:14:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Mars 2020 ascent rate Maria Zuber +2025-04-04 at 04:14:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +boost phases were within acceptable spacecraft structural. design limits. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:14:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 descent propulsion system start time April 11 1970" +2025-04-04 at 04:14:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:14:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:14:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Mars 2020 ascent trajectory lift-off to descent engine firing +2025-04-04 at 04:14:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:14:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:14:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Mars 2020 ascent altitude rate +2025-04-04 at 04:14:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-04 at 04:14:14 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:14:14 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:14:14 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, True, False, True] +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1771, 520, 514, 1172, 278, 393] +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_correctness:84 - Average student length: 774.67 +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_correctness:86 - Length ratio: 77.47 +2025-04-04 at 04:14:14 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-04 at 04:14:14 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.400 Âą 0.409 +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.33 Âą 2.62 +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 7, 3, 0, 4] +2025-04-04 at 04:14:14 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +boost phases were within acceptable spacecraft structural. design limits. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +boost phases were within acceptable spacecraft structural. design limits. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-04 at 04:14:14 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nboost phases were within acceptable spacecraft structural. design limits.\n---...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nTABLE 4-II.- TRAJECTORY PARAMETERS\n------\nResult 2:\n1.0 SUMMARY 1-1 2.0 INTRO...', 'Result 1:\nIn this report, all actual times prior to earth landing are elapsed time from...', 'Result 1:\nThe analysis of the trajectory from lift-off to spacecraft/S-IVB separation w...', 'Result 1:\nboost phases were within acceptable spacecraft structural. design limits.\n---...', 'Result 1:\nThe vehicle was launched on an azimuth 90 degrees east of north, and a roll m...', 'Result 1:\nIn this report, all actual times prior to earth landing are elapsed time from...'] +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:384 - īŋŊīŋŊ Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:14:14 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nPowering down of the command and service nodules and powering up of the lunar...'] +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:14:14 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nEvent Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-o...'] +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:14:14 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nIn this report, all actual times prior to earth landing are elapsed time from...', 'Result 1:\nIn this report, all actual times prior to earth landing are elapsed time from...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...'] +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:14:14 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...'] +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-04 at 04:14:14 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 0.0 +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.000 +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.000 +2025-04-04 at 04:14:14 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.433, Perfect scores: 2/6 +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 2.83 +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:14:14 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:14:14 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.426, Max reward: 0.669 +2025-04-04 at 04:14:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:14:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module carbon dioxide removal system +2025-04-04 at 04:14:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module carbon dioxide removal system +2025-04-04 at 04:14:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo lunar module eclss carbon dioxide scrubber +2025-04-04 at 04:14:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-04 at 04:14:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: carbon dioxide removal system lunar module Apollo +2025-04-04 at 04:14:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: carbon dioxide removal system lunar module Apollo +2025-04-04 at 04:14:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:14:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module lithium hydroxide carbon dioxide removal system +2025-04-04 at 04:14:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-04 at 04:14:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module carbon dioxide removal system command module lithium hydroxide cartridges +2025-04-04 at 04:14:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-04 at 04:14:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo lunar module carbon dioxide scrubbing system name +2025-04-04 at 04:14:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-04 at 04:14:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module carbon dioxide removal system command module lithium hydroxide +2025-04-04 at 04:14:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-04 at 04:14:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module lunar module environmental control system carbon dioxide removal procedure +2025-04-04 at 04:14:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-04 at 04:14:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:14:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module carbon dioxide scrubbing system +2025-04-04 at 04:14:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-04 at 04:14:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module carbon dioxide system improvisation commande module cartridges +2025-04-04 at 04:14:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: advanced life support system apollo lunar module +2025-04-04 at 04:14:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:14:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module carbon dioxide removal systems procedure +2025-04-04 at 04:14:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-04 at 04:14:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:14:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module carbon dioxide carbon capture +2025-04-04 at 04:14:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module carbon dioxide removal system setup Apollo +2025-04-04 at 04:14:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system air purification +2025-04-04 at 04:14:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lithium hydroxide cartridges lunar module carbon dioxide removal procedure +2025-04-04 at 04:14:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-04 at 04:14:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:14:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar module carbon dioxide removal system emergency procedure +2025-04-04 at 04:14:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:14:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lithium hydroxide cartridge lunar module carbon dioxide removal effectiveness +2025-04-04 at 04:14:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-04 at 04:14:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:14:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module carbon dioxide removal trade-offs +2025-04-04 at 04:14:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:14:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module carbon dioxide removal integrated system +2025-04-04 at 04:14:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:14:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module life support system overview +2025-04-04 at 04:14:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-04 at 04:14:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:14:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module thermal control system reliability +2025-04-04 at 04:14:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:14:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:14:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module docking tunnel insulation issue +2025-04-04 at 04:14:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-04 at 04:14:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:14:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module power system changes +2025-04-04 at 04:14:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-04 at 04:14:38 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:14:38 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:14:38 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, True, True, False, False] +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_correctness:82 - Student lengths: [211, 370, 591, 482, 1772, 597] +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [28, 28, 28, 28, 28, 28] +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_correctness:84 - Average student length: 670.50 +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 28.00 +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_correctness:86 - Length ratio: 23.95 +2025-04-04 at 04:14:38 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-04 at 04:14:38 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.350 Âą 0.364 +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 1.67 Âą 1.80 +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [4, 0, 4, 2, 0, 0] +2025-04-04 at 04:14:38 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 12: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-04 at 04:14:38 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 5.0 +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.833 +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:14:38 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.267, Perfect scores: 1/6 +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 0.67 +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:14:38 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:14:38 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.357, Max reward: 0.620 +2025-04-04 at 04:14:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:14:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is S-IC boost and how does it relate to maximum lateral loads compared to design value? +2025-04-04 at 04:14:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What percentage of maximum lateral loads experienced during S-IC boost of Tesla's Semi compared to design value +2025-04-04 at 04:14:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the percentage of lateral load experienced during S-IC boost in Nissan GT-R +2025-04-04 at 04:14:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the percentage of maximum lateral loads during S-IC boost compared to design value +2025-04-04 at 04:14:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the percentage of maximum lateral loads during S-IC boost in electric vehicles +2025-04-04 at 04:14:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What percentage of maximum lateral loads experienced during S-IC boost compared to design value +2025-04-04 at 04:14:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:14:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IC engine boost characteristics and maximum lateral loads +2025-04-04 at 04:14:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what percent of max lateral loads during S-IC boost exceeds design value for Tesla Semi +2025-04-04 at 04:14:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the percentage of maximum lateral load experienced during S-IC boost compared to design value +2025-04-04 at 04:14:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the design value of maximum lateral loads for S-IC boost in the S-II mission? +2025-04-04 at 04:14:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-IC boost maximum lateral load percentage design value +2025-04-04 at 04:14:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:14:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission S-IC boost peak acceleration and oscillation data +2025-04-04 at 04:14:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-04 at 04:14:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the maximum lateral load experienced during S-IC boost of Saturn S-II +2025-04-04 at 04:14:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:14:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the design value for maximum lateral load during S-IC boost of Saturn S-II +2025-04-04 at 04:14:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:14:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the design load limit for structural loads during S-IC boost of Saturn S-II +2025-04-04 at 04:14:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:14:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the maximum lateral load used to derive 25% in Result 1 +2025-04-04 at 04:14:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +boost phases were within acceptable spacecraft structural. design limits. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:14:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:14:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the maximum lateral load experienced during S-IC boost of Saturn S-II +2025-04-04 at 04:14:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:14:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the design value of maximum lateral load experienced during S-IC with peak 33.7g +2025-04-04 at 04:14:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:58 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:14:58 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:14:58 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, False, True, True, False] +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.67 +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_correctness:82 - Student lengths: [422, 250, 2047, 245, 670, 17] +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_correctness:84 - Average student length: 608.50 +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_correctness:86 - Length ratio: 60.85 +2025-04-04 at 04:14:58 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:14:58 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.517 Âą 0.281 +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.83 Âą 2.48 +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [3, 2, 8, 2, 2, 0] +2025-04-04 at 04:14:58 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +boost phases were within acceptable spacecraft structural. design limits. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 6.0 +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 1.000 +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.000 +2025-04-04 at 04:14:58 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.533, Perfect scores: 0/6 +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 0.67 +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 4/6 +2025-04-04 at 04:14:58 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:14:58 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.570, Max reward: 0.732 +2025-04-04 at 04:15:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:15:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the factors airlines consider when setting flight menus? airlines menu selection process? +2025-04-04 at 04:15:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-04 at 04:15:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: meal selection algorithm +2025-04-04 at 04:15:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-04 at 04:15:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Why airlines choose flight options" or "How do airlines select flight menus" +2025-04-04 at 04:15:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-04 at 04:15:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: airline flight menu selection basis +2025-04-04 at 04:15:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-04 at 04:15:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the basis for airline menu selection? +2025-04-04 at 04:15:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-04 at 04:15:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:15:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the specific consideration for selecting flight menus in emergency situations? +2025-04-04 at 04:15:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-04 at 04:15:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: selecting food and water for spaceflight +2025-04-04 at 04:15:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-04 at 04:15:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Spaceflight food selection criteria" or "Spaceflight menu planning" +2025-04-04 at 04:15:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-04 at 04:15:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: factors influencing airline flight menu selection +2025-04-04 at 04:15:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-04 at 04:15:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 crew menu selection Apollo spaceflight +2025-04-04 at 04:15:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:15:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:15:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What selection criteria are used for flight menus? +2025-04-04 at 04:15:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-04 at 04:15:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: food packaging and storage in spaceflight +2025-04-04 at 04:15:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:15:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Spaceflight food menu planning process" or "Crew-preference based food selection" +2025-04-04 at 04:15:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-04 at 04:15:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: selection of flight menus based on factors like water conservation and crew preferences +2025-04-04 at 04:15:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-04 at 04:15:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 mission crew training objectives +2025-04-04 at 04:15:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-04 at 04:15:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:15:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA checklist for mission success +2025-04-04 at 04:15:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:15:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: crew preferences for flight menu selection +2025-04-04 at 04:15:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-04 at 04:15:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:15:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft evaluation assessments +2025-04-04 at 04:15:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:15:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollos missions had preflight evaluation of crew preferences +2025-04-04 at 04:15:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:15:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:15:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ashurs testing and inspection +2025-04-04 at 04:15:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:15:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission performance and analysis +2025-04-04 at 04:15:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:15:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:15:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: operation procedures after apollo 8 +2025-04-04 at 04:15:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:15:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo crew training for emergency response +2025-04-04 at 04:15:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:15:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:15:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: optimal lunar landing strategy +2025-04-04 at 04:15:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-04 at 04:15:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 training schedule and impacts +2025-04-04 at 04:15:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:15:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:15:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: coordination challenges of training for Apollo 13 +2025-04-04 at 04:15:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:15:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:15:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 training for solar alignment and debris avoidance +2025-04-04 at 04:15:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:15:27 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:15:27 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:15:27 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, True, False, False] +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_correctness:82 - Student lengths: [767, 1568, 614, 799, 2044, 321] +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [82, 82, 82, 82, 82, 82] +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_correctness:84 - Average student length: 1018.83 +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 82.00 +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_correctness:86 - Length ratio: 12.42 +2025-04-04 at 04:15:27 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:15:27 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.188 Âą 0.270 +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 6.67 Âą 13.61 +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 37, 0, 3, 0, 0] +2025-04-04 at 04:15:27 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-04 at 04:15:27 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe crew followed the flight menus prior to the inflight incident and maintai...', 'Result 1:\nThe crew followed the flight menus prior to the inflight incident and maintai...', 'Result 1:\nThe crew followed the flight menus prior to the inflight incident and maintai...'] +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-04 at 04:15:27 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:15:27 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.367, Perfect scores: 1/6 +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 2.50 +2025-04-04 at 04:15:27 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:15:27 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:15:28 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.467, Max reward: 0.652 +2025-04-04 at 04:15:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:15:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: shaft rotation zero output equivalent +2025-04-04 at 04:15:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-04 at 04:15:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: and +2025-04-04 at 04:15:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +(section ll.3). +------ + +2025-04-04 at 04:15:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "no-load state" +2025-04-04 at 04:15:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +venting of the helium tank would be cause for aborting the mission, the unwanted rolling moment, which is quite small, would have no ultimate effect on a nominal profile. Therefore, the vent tube configuration for future spacecraft will not be changed to one having zero net thrust. +------ + +2025-04-04 at 04:15:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: equivalent of zero output in shaft rotation +2025-04-04 at 04:15:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-04 at 04:15:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:15:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: shaft rotation zero optics mode +2025-04-04 at 04:15:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:15:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: and +2025-04-04 at 04:15:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +(section ll.3). +------ + +2025-04-04 at 04:15:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "no-load rotation" +2025-04-04 at 04:15:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +venting of the helium tank would be cause for aborting the mission, the unwanted rolling moment, which is quite small, would have no ultimate effect on a nominal profile. Therefore, the vent tube configuration for future spacecraft will not be changed to one having zero net thrust. +------ + +2025-04-04 at 04:15:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:15:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "zero net thrust" +2025-04-04 at 04:15:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +venting of the helium tank would be cause for aborting the mission, the unwanted rolling moment, which is quite small, would have no ultimate effect on a nominal profile. Therefore, the vent tube configuration for future spacecraft will not be changed to one having zero net thrust. +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:15:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:15:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "zero net moment" +2025-04-04 at 04:15:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:15:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:15:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "mean product of inertia" +2025-04-04 at 04:15:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:15:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:15:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "mean thrust" +2025-04-04 at 04:15:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-04 at 04:15:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:15:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "mean velocity translation" +2025-04-04 at 04:15:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-04 at 04:15:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:15:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "mean rate of change" +2025-04-04 at 04:15:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +-1.2 40.4 +0.4 +------ +Result 2: +The potable water quantity measurement fluctuated briefly on two occasions during the mission. At about 23 hours, the reading decreased from 98 to 79 percent for about 5 minutes and then returned to a normal reading of approximately l02 percent. Another fluctuation was noted at about 37 hours, at which time the reading decreased from its upper limit to 83.5 percent. The reading then returned to the upper limit in a period of 7 seconds. + +Preflight fluctuations of from 2 to 6 percent near the full level were observed once during the countdown demonstration test, and a possible earlier fluctuation of about 4 percent at the half-full level was noted during the flight readiness test. +------ + +2025-04-04 at 04:15:44 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:15:44 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:15:44 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_correctness:82 - Student lengths: [407, 175, 371, 764, 659, 683] +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_correctness:84 - Average student length: 509.83 +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 12.00 +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_correctness:86 - Length ratio: 42.49 +2025-04-04 at 04:15:44 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-04 at 04:15:44 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.079 Âą 0.177 +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 4.17 Âą 9.32 +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 0, 25, 0, 0] +2025-04-04 at 04:15:44 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: ++21.0 +3.0 +------ +Result 2: +(section ll.3). +------ + +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: ++21.0 +3.0 +------ +Result 2: +(section ll.3). +------ + +2025-04-04 at 04:15:44 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\n+21.0 +3.0\n------\nResult 2:\n(section ll.3).\n------\n...', 'Result 1:\n+21.0 +3.0\n------\nResult 2:\n(section ll.3).\n------\n...'] +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-04 at 04:15:44 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +venting of the helium tank would be cause for aborting the mission, the unwanted rolling moment, which is quite small, would have no ultimate effect on a nominal profile. Therefore, the vent tube configuration for future spacecraft will not be changed to one having zero net thrust. +------ + +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +venting of the helium tank would be cause for aborting the mission, the unwanted rolling moment, which is quite small, would have no ultimate effect on a nominal profile. Therefore, the vent tube configuration for future spacecraft will not be changed to one having zero net thrust. +------ + +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +venting of the helium tank would be cause for aborting the mission, the unwanted rolling moment, which is quite small, would have no ultimate effect on a nominal profile. Therefore, the vent tube configuration for future spacecraft will not be changed to one having zero net thrust. +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +-1.2 40.4 +0.4 +------ +Result 2: +The potable water quantity measurement fluctuated briefly on two occasions during the mission. At about 23 hours, the reading decreased from 98 to 79 percent for about 5 minutes and then returned to a normal reading of approximately l02 percent. Another fluctuation was noted at about 37 hours, at which time the reading decreased from its upper limit to 83.5 percent. The reading then returned to the upper limit in a period of 7 seconds. + +Preflight fluctuations of from 2 to 6 percent near the full level were observed once during the countdown demonstration test, and a possible earlier fluctuation of about 4 percent at the half-full level was noted during the flight readiness test. +------ + +2025-04-04 at 04:15:44 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nEfforts to install the tunnel hatch were terminated when the Commander observ...', 'Result 1:\nThe passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not ...', 'Result 1:\nventing of the helium tank would be cause for aborting the mission, the unwan...', 'Result 1:\nEvent Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product o...', 'Result 1:\nEvent Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product o...', 'Result 1:\nEvent Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product o...', 'Result 1:\nPlatform-sensed velocity changes, ft/sec Command module axes Lunar module axe...', 'Result 1:\n-1.2 40.4 +0.4\n------\nResult 2:\nThe potable water quantity measurement fluctu...'] +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-04 at 04:15:44 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 2.0 +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.333 +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:15:44 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.067, Perfect scores: 0/6 +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 1/6 +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 0.17 +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 0/6 +2025-04-04 at 04:15:44 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:15:44 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.120, Max reward: 0.518 +2025-04-04 at 04:15:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:15:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: publication date of report titled Performance Analysis +2025-04-04 at 04:15:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:15:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: publication date of "Performance Analysis report" +2025-04-04 at 04:15:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:15:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: publication date of Performance Analysis report +2025-04-04 at 04:15:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:15:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: report titled Performance Analysis publication date finance +2025-04-04 at 04:15:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:15:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: publication date of the report titled Performance Analysis +2025-04-04 at 04:15:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:15:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Report "Performance Analysis" publication date +2025-04-04 at 04:15:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:15:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:15:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: publication date Apollo 7 through Apollo 13 mission reports +2025-04-04 at 04:15:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:15:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: report titled Performance Analysis publication date NASA +2025-04-04 at 04:15:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:15:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 7-13 mission report Performance Analysis table E-I +2025-04-04 at 04:15:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:15:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Performance Analysis report publication date engineering +2025-04-04 at 04:15:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:15:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo performance analysis report publication date +2025-04-04 at 04:15:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:15:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:15:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 7 Apollo 9 Performance Analysis +2025-04-04 at 04:15:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:15:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Performance Analysis report publication date +2025-04-04 at 04:15:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:15:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 7-13 Performance Analysis report +2025-04-04 at 04:15:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:15:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Performance Analysis report publication date by organization such as IEEE +2025-04-04 at 04:15:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:15:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 Performance Analysis report publication date +2025-04-04 at 04:15:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:15:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:15:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo report titled Performance Analysis +2025-04-04 at 04:15:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:15:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Command and Service Module Performance Analysis report +2025-04-04 at 04:15:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:15:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Performance Analysis report published in 1970 +2025-04-04 at 04:15:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:15:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 Performance Analysis report +2025-04-04 at 04:15:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:15:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:15:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo report Performance Analysis by date +2025-04-04 at 04:15:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:15:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 7 postflight performance analysis report +2025-04-04 at 04:15:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:15:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Performance of Lunar Module Reaction Control System Review publication date +2025-04-04 at 04:15:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:15:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:16:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: report Performance Summary Apollo 7 +2025-04-04 at 04:16:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:16:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 7 postflight report supplement +2025-04-04 at 04:16:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:16:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Reaction Control System Performance Analysis report December 1969 to September 1970 +2025-04-04 at 04:16:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:16:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:16:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 Supplement 7 Performance Analysis report +2025-04-04 at 04:16:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:16:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Performance Analysis report publication date and title +2025-04-04 at 04:16:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:16:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:16:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Supplement 5 Performance Analysis +2025-04-04 at 04:16:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:16:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Performance Analysis December 1969 +2025-04-04 at 04:16:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:16:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:16:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Performance Analysis postflight report +2025-04-04 at 04:16:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:16:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Performance Analysis report December 1969 +2025-04-04 at 04:16:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:16:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:16:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module Supplement 7 postflight analysis +2025-04-04 at 04:16:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:16:11 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:16:11 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:16:11 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, False, False, False, False] +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_correctness:82 - Student lengths: [144, 427, 1397, 267, 197, 346] +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_correctness:84 - Average student length: 463.00 +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 13.00 +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_correctness:86 - Length ratio: 35.62 +2025-04-04 at 04:16:11 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_format:228 - Responses ending properly: 6/6 +2025-04-04 at 04:16:11 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.400 Âą 0.409 +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.17 Âą 2.34 +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [3, 6, 0, 0, 4, 0] +2025-04-04 at 04:16:11 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:16:11 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...'] +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:16:11 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...'] +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:16:11 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...', 'Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...', 'Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...', 'Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...'] +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:16:11 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...'] +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:16:11 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nTable E-I contains a listing of all supplemental reports that are or will be ...'] +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 1.0 +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.167 +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:16:11 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.467, Perfect scores: 1/6 +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 5/6 +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 3.50 +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:16:11 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:16:11 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.420, Max reward: 0.574 +2025-04-04 at 04:16:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:16:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the date of the document known as Mission Evaluation Team Report that was possibly approved in 2023 or 2022 together with NASA +2025-04-04 at 04:16:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:16:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: WHO document approval year Mission Evaluation Team +2025-04-04 at 04:16:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The Department of Defense provided recovery support commensurate with mission planning for Apollo l3. Because of the emergency which resulted in premature termination of the mission, additional support was provided by the Department of Defense and offers of assistance were made by many foreign nations, including England, France, Greece, Spain, Germany, Uruguay, Brazil, Kenya, the Netherlands, Nationalist China, and the Soviet Union. As a result of this voluntary support, a total of 2l ships and 17 aircraft were available for supporting an Indian Ocean landing, and 5l ships and 2l aircraft for an Atlantic Ocean landing. In the Pacific Ocean, there were l3 ships and 17 aircraft known to be available over and above the forces designated for primary recovery support. +------ + +2025-04-04 at 04:16:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what year was the document approved by the Mission Evaluation Team in reference to the Mars sample return mission +2025-04-04 at 04:16:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:16:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Artemis Mission Evaluation Team document approval year +2025-04-04 at 04:16:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:16:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Mission Evaluation document approval year +2025-04-04 at 04:16:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:16:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:16:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: supplemental reports for apollo missions published 1970 +2025-04-04 at 04:16:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:16:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Mission Evaluation Team approved report +2025-04-04 at 04:16:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:16:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Mission Evaluation Team document approval date +2025-04-04 at 04:16:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:16:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Mission Evaluation Team document approval year +2025-04-04 at 04:16:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:16:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:16:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: document approved by mission evaluation team september 1970 +2025-04-04 at 04:16:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:16:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Mission Evaluation Team approved document +2025-04-04 at 04:16:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:16:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Mission Briefing Document +2025-04-04 at 04:16:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:16:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Mission Evaluation Team meeting minutes or report approval date +2025-04-04 at 04:16:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:16:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:16:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 Mission Evaluation Team Report document +2025-04-04 at 04:16:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:16:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:24 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:16:24 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:16:24 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, False, False, True, True] +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.67 +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_correctness:82 - Student lengths: [451, 296, 124, 605, 217, 633] +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_correctness:84 - Average student length: 387.67 +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 4.00 +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_correctness:86 - Length ratio: 96.92 +2025-04-04 at 04:16:24 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-04 at 04:16:24 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.404 Âą 0.292 +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.83 Âą 2.67 +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [3, 3, 0, 0, 8, 3] +2025-04-04 at 04:16:24 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The Department of Defense provided recovery support commensurate with mission planning for Apollo l3. Because of the emergency which resulted in premature termination of the mission, additional support was provided by the Department of Defense and offers of assistance were made by many foreign nations, including England, France, Greece, Spain, Germany, Uruguay, Brazil, Kenya, the Netherlands, Nationalist China, and the Soviet Union. As a result of this voluntary support, a total of 2l ships and 17 aircraft were available for supporting an Indian Ocean landing, and 5l ships and 2l aircraft for an Atlantic Ocean landing. In the Pacific Ocean, there were l3 ships and 17 aircraft known to be available over and above the forces designated for primary recovery support. +------ + +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-04 at 04:16:24 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:16:24 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...'] +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:16:24 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.167, Perfect scores: 1/6 +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 1/6 +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 0.33 +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:16:24 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:16:24 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.341, Max reward: 0.642 +2025-04-04 at 04:16:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:16:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the location of the NASA Johnson Space Center +2025-04-04 at 04:16:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa manned space center location +2025-04-04 at 04:16:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Lyndon B. Johnson Space Center +2025-04-04 at 04:16:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa manned space shuttle center location +2025-04-04 at 04:16:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the location of the NASA Johnson Space Center +2025-04-04 at 04:16:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:16:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: address of NASA Johnson Space Center +2025-04-04 at 04:16:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of NASA-Johnson Space Center +2025-04-04 at 04:16:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:16:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Johnson Space Center Houston Texas address +2025-04-04 at 04:16:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: google maps nasa johson space center houston tx +2025-04-04 at 04:16:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:16:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: current address of nasa johson space center houston tx +2025-04-04 at 04:16:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa johson space center current address 2023 +2025-04-04 at 04:16:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa johson space center houston texas contact address +2025-04-04 at 04:16:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa houston address map +2025-04-04 at 04:16:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:16:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa johson space center 2023 contact and address +2025-04-04 at 04:16:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: why is nasa houston address not public +2025-04-04 at 04:16:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa johson space center houston why publicly available +2025-04-04 at 04:16:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:16:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa jhonn space center agency contact person +2025-04-04 at 04:16:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-04 at 04:16:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa jhonn space center contact information +2025-04-04 at 04:16:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: google nasa houston contact person +2025-04-04 at 04:16:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +The first reported electronic contact by the recovery forces was through S-band contact by Samoa Rescue 4. A visual sighting report by the Recovery helicopter was received and was followed shortly thereafter by aquisition of the recovery beacon signal by the Recovery, Photo, and Swim l helicopters. Fuel dump was noted and voice contact was made with the descending spacecraft, although no latitude and longitude data were received. The command module landed at 1807 G.m.t. and remained in the stable l flotation attitude. The flashing light was operating and the infiation of the uprighting system commenced about l0 minutes subsequent to landing. +------ + +2025-04-04 at 04:16:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa jhonn space center contact information email +2025-04-04 at 04:16:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa johson space center phone number +2025-04-04 at 04:16:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa jhonn space center address map +2025-04-04 at 04:16:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa jhonn space center location +2025-04-04 at 04:16:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:16:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:16:45 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:16:45 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:16:45 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, True, True, True, False] +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.83 +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_correctness:82 - Student lengths: [168, 304, 285, 120, 174, 121] +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [20, 20, 20, 20, 20, 20] +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_correctness:84 - Average student length: 195.33 +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 20.00 +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_correctness:86 - Length ratio: 9.77 +2025-04-04 at 04:16:45 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:16:45 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.417 Âą 0.282 +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 3.83 Âą 6.36 +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [1, 18, 1, 2, 1, 0] +2025-04-04 at 04:16:45 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 12: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 13: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 14: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +The first reported electronic contact by the recovery forces was through S-band contact by Samoa Rescue 4. A visual sighting report by the Recovery helicopter was received and was followed shortly thereafter by aquisition of the recovery beacon signal by the Recovery, Photo, and Swim l helicopters. Fuel dump was noted and voice contact was made with the descending spacecraft, although no latitude and longitude data were received. The command module landed at 1807 G.m.t. and remained in the stable l flotation attitude. The flashing light was operating and the infiation of the uprighting system commenced about l0 minutes subsequent to landing. +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 15: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 16: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 17: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 18: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...'] +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-04 at 04:16:45 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:16:45 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.233, Perfect scores: 1/6 +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.00 +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:16:45 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:16:45 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.265, Max reward: 0.720 +2025-04-04 at 04:16:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:16:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Did the pilot mention a function during a mission in a NASA press conference or a specific incident involving a pilot and a function? +2025-04-04 at 04:16:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-04 at 04:16:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Was pilot involved in describing a function during Apollo mission +2025-04-04 at 04:16:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ + +2025-04-04 at 04:16:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 oxygen tank accident oral protocol" +2025-04-04 at 04:16:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-04 at 04:16:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Was pilot involved in describing function during a specific space mission or spacecraft operation? +2025-04-04 at 04:16:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-04 at 04:16:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: was pilot involved in describing a function during a space mission +2025-04-04 at 04:16:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-04 at 04:16:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:16:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Details of Apollo 13 mission about whether a pilot's function was described during the mission +2025-04-04 at 04:16:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:16:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission pilot involvment 55 hour 54 minute event +2025-04-04 at 04:16:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:16:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 mission flight controllers communication" +2025-04-04 at 04:16:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:16:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Was a pilot involved in manually controlling the spacecraft during a specific experiment or anomaly on the Apollo 13 mission? +2025-04-04 at 04:16:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:16:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what happened during the loud noise incident in apollo 13 that resulted in a master alarm and master bus undervoltage condition +2025-04-04 at 04:16:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:16:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:16:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 pilot described function during mission +2025-04-04 at 04:16:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-04 at 04:16:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission re-entry pilot +2025-04-04 at 04:16:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:16:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 mission crew communication with Mission Control" +2025-04-04 at 04:16:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:16:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What role did the lunar module pilot play during the Cryogenic Oxygen Tank Incident on the Apollo 13 mission? +2025-04-04 at 04:16:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ + +2025-04-04 at 04:16:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 master alarm and master bus undervoltage condition overview +2025-04-04 at 04:16:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:16:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:16:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 pilot during experiment T-029 +2025-04-04 at 04:16:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:16:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module pilot Apollo 8 navigation +2025-04-04 at 04:16:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:16:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 mission pilot communication during lunar landing" +2025-04-04 at 04:16:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:16:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 thumping noise surge and electrical power system issues +2025-04-04 at 04:16:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:16:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:17:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module pilot experiment T-029 +2025-04-04 at 04:17:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-04 at 04:17:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 pilot did not involve in first transearth midcourse correction +2025-04-04 at 04:17:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:17:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 13 pilot talking points jims Lovell" +2025-04-04 at 04:17:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-04 at 04:17:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar module descent stage electrical system short circuit +2025-04-04 at 04:17:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:17:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:17:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 pilot involvement in transearth correction +2025-04-04 at 04:17:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:17:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot's function during apollo 13 mission involving thumping noise and current surge +2025-04-04 at 04:17:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-04 at 04:17:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:17:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth correction performed by crew or automated system +2025-04-04 at 04:17:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-04 at 04:17:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 pilot's role during descent stage noise and current spikes +2025-04-04 at 04:17:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:17:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:17:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection performed by lunar module reaction control system +2025-04-04 at 04:17:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-04 at 04:17:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot's description of lunar module descent stage short circuit +2025-04-04 at 04:17:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:17:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:17:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: was the pilot involved in lunar module reaction control +2025-04-04 at 04:17:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:17:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: function during apollo 13 mission involving lunar module ascent projectile command +2025-04-04 at 04:17:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:17:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:17:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 Reaction control system manned operations +2025-04-04 at 04:17:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:17:10 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:17:10 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:17:10 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, True, False, True, False] +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_correctness:82 - Student lengths: [477, 1841, 601, 532, 586, 1986] +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [3, 3, 3, 3, 3, 3] +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_correctness:84 - Average student length: 1003.83 +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 3.00 +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_correctness:86 - Length ratio: 334.61 +2025-04-04 at 04:17:10 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.833, Valid formats: 5.0/6 +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:17:10 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.425 Âą 0.437 +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.17 Âą 2.27 +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [5, 0, 5, 0, 3, 0] +2025-04-04 at 04:17:10 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-04 at 04:17:10 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe television presentation during the midcourse correction maneuver, as well...', 'Result 1:\nThe pilot describing function experiment (T-029) was a success, in that data ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:17:10 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...'] +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-04 at 04:17:10 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ + +2025-04-04 at 04:17:10 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe television presentation during the midcourse correction maneuver, as well...', 'Result 1:\nThe pilot describing function experiment (T-029) was a success, in that data ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:17:10 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe television presentation during the midcourse correction maneuver, as well...', 'Result 1:\nThe field-change and sferics detectors at site 5 gave no indication of any li...', 'Result 1:\nApproximately 2 seconds later, the Command Module Pilot reported a master ala...', 'Result 1:\nAt 97 hours 14 minutes, the crew reported a thumping noise and snowflakes ven...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nAt 97 hours 14 minutes, the crew reported a thumping noise and snowflakes ven...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 1.0 +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.167 +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:17:10 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.567, Perfect scores: 2/6 +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 5/6 +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 3.67 +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:17:10 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:17:10 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.447, Max reward: 0.715 +2025-04-04 at 04:17:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:17:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How much cooling hours of 50 pounds of water at reduced power condition in a typical system +2025-04-04 at 04:17:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Specific heat capacity of 50 pounds of water, conversion to British Thermal Units (BTU) and hours +2025-04-04 at 04:17:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Can there be a search query that says: "water cooling power reduction rate?" or "ways to reduce power for cooling systems"? +2025-04-04 at 04:17:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:17:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 50 pounds of water heat sink cooling efficiency +2025-04-04 at 04:17:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +------ + +2025-04-04 at 04:17:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the thermal conductivity of water at 20°C, and how does it relate to cooling efficiency? +2025-04-04 at 04:17:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:17:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:17:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How many hours of cooling can 50 pounds of water provide at 350-400 watts, as in the space shuttle, at cabin temperatures around 55 degrees Fahrenheit +2025-04-04 at 04:17:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Specific heat capacity of water, specific heat capacity of potable water in pounds of water and temperature reduction from 54-60 degrees Fahrenheit to 43 degrees Fahrenheit +2025-04-04 at 04:17:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:17:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how much cooling water provides at 350-400 watts +2025-04-04 at 04:17:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-04 at 04:17:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft cabin cooling reduced power condition water consumption +2025-04-04 at 04:17:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:17:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How does the thermal conductivity of water reduce with a reduction in power? +2025-04-04 at 04:17:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:17:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:17:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water cooling system capacity, 50 pounds of water, 350-400 watts, cooling time +2025-04-04 at 04:17:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-04 at 04:17:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Specific heat capacity of water and conversion from BTU to the cooling of a cabin with temperature of 50-60°F to 43°F +2025-04-04 at 04:17:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how much water can provide 5 hours of cooling at 350-400 watts +2025-04-04 at 04:17:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:17:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft cabin temperature reduction 50 pounds water cooling capacity +2025-04-04 at 04:17:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How much cooling can 1 liter of water provide at 350-400 watts power? +2025-04-04 at 04:17:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:17:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:17:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: hybrid cooling system capacity, power consumption rate, 50 pounds of water, water cooling system cooling time, space shuttle +2025-04-04 at 04:17:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-04 at 04:17:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Specific heat capacity of water, coefficient of cooling, 50°F - 60°F reduction, BTU +2025-04-04 at 04:17:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how much cooling provided by 50 pounds of water under 350-400 watts +2025-04-04 at 04:17:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-04 at 04:17:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module cooling water tank capacity 50 pounds +2025-04-04 at 04:17:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How does the amount of cooling provided by 1 liter of water change at different power levels, such as 10% or 50% of maximum power? +2025-04-04 at 04:17:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:17:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:17:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water separator drain tank water flow rate, 50 pounds of water, water flow rate through water separator +2025-04-04 at 04:17:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-04 at 04:17:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 126,789 BTUs to hours, BTU per hour to pounds of water +2025-04-04 at 04:17:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:17:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how much water cooling provided by 50 pounds of water for 5 hours +2025-04-04 at 04:17:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:17:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module suit compressor water usage +2025-04-04 at 04:17:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-04 at 04:17:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Cooling efficiency of a liquid (water) and how it relates to power reduction in a general context +2025-04-04 at 04:17:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:17:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: oxygen system water usage, water separator operation, 50 pounds of water, cabin temperature +2025-04-04 at 04:17:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-04 at 04:17:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pounds of water used from the lunar module tanks between activation of the sublimator and undocking, 50 pounds of water remaining +2025-04-04 at 04:17:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-04 at 04:17:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water usage rate between 2.6-6.3 lb/hr +2025-04-04 at 04:17:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-04 at 04:17:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: oxygen usage water consumption command module 50 pounds +2025-04-04 at 04:17:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-04 at 04:17:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Cooling efficiency of a liquid for a given surface area or volume and a power reduction factor +2025-04-04 at 04:17:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:17:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:17:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: potable water tank water flow rate, oxygen system water use rate, 50 pounds of water +2025-04-04 at 04:17:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-04 at 04:17:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 52 pounds of water remaining, 24 pounds of water used +2025-04-04 at 04:17:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-04 at 04:17:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water usage rate 2.6-6.3 lb/hr +2025-04-04 at 04:17:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-04 at 04:17:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: oxygen breathing rate water cooling capacity 50 pounds +2025-04-04 at 04:17:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-04 at 04:17:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Thermal conductivity of water and its relation to mass, surface area, or volume for cooling +2025-04-04 at 04:17:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:17:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water cooling system, potable water tank water usage, oxygen system water supply +2025-04-04 at 04:17:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-04 at 04:17:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: consumed water lunar module tank, hourly rate +2025-04-04 at 04:17:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:17:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: average of 2.6-6.3 lb/hr +2025-04-04 at 04:17:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ +Result 2: +-1.2 40.4 +0.4 +------ + +2025-04-04 at 04:17:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module water cooling capacity +2025-04-04 at 04:17:37 | INFO | src.agent:run_tool_calls:118 - īŋŊīŋŊī¸ Information: Result 1: +arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:17:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Cooling efficiency of a liquid with respect to its mass, assuming a fixed volume or surface area +2025-04-04 at 04:17:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:17:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:17:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ash, contamination, oxygen system, water cooling system +2025-04-04 at 04:17:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-04 at 04:17:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: metabolic consumption rate 0.2101b/hr +2025-04-04 at 04:17:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +-1.2 40.4 +0.4 +------ + +2025-04-04 at 04:17:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Cooling rate of liquid (water) per unit mass, assuming a given volume or surface area +2025-04-04 at 04:17:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:17:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:17:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water from lunar module tanks used in 5 hours +2025-04-04 at 04:17:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-04 at 04:17:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Specific heat capacity of water and its relation to mass, assuming a fixed volume or surface area +2025-04-04 at 04:17:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-04 at 04:17:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:17:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Cooling efficiency of a liquid (water) for evaporation or boiling, assuming a fixed volume or surface area +2025-04-04 at 04:17:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:17:44 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:17:44 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:17:44 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 1/6 answers correct +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, True, False] +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1692, 1544, 1997, 421, 218, 1009] +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_correctness:84 - Average student length: 1146.83 +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 8.00 +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_correctness:86 - Length ratio: 143.35 +2025-04-04 at 04:17:44 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.833, Valid formats: 5.0/6 +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:17:44 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.079 Âą 0.177 +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.00 Âą 4.47 +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 0, 0, 12, 0] +2025-04-04 at 04:17:44 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-04 at 04:17:44 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nHy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 ...', 'Result 1:\nThe oxygen distribution system operated nominally until deactivation followin...', 'Result 1:\nHy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 ...', 'Result 1:\nThe oxygen distribution system operated nominally until deactivation followin...', 'Result 1:\nASHUR Purpose Tests performed Results Environmental Control 109007 To determi...'] +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:17:44 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nDuring the period when the command module was powered down, the cabin tempera...', 'Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nHy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 ...', 'Result 1:\nHy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 ...', 'Result 1:\nHy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 ...', 'Result 1:\nTotal oxygen usage from the three lunar module oxygen tanks was 20.3 pounds o...'] +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ +Result 2: +-1.2 40.4 +0.4 +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +-1.2 40.4 +0.4 +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-04 at 04:17:44 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\npropulsion system was changed to include a bypass line around the fuel/helium...', 'Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nHy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 ...', 'Result 1:\n6.7 ENVIRONMENTAL CONTROL\n\nEnvironmental. control system performance was sati...', 'Result 1:\n6.7 ENVIRONMENTAL CONTROL\n\nEnvironmental. control system performance was sati...', 'Result 1:\nTotal oxygen usage from the three lunar module oxygen tanks was 20.3 pounds o...', 'Result 1:\nFuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5...', 'Result 1:\n6.7 ENVIRONMENTAL CONTROL\n\nEnvironmental. control system performance was sati...'] +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-04 at 04:17:44 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:17:44 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe most likely cause of the anomaly is a tank-insulation degradation which w...', 'Result 1:\nDuring the period when the command module was powered down, the cabin tempera...', 'Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nElectrical shorts in the fan circuit ignited the wire insulation, causing pre...', 'Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nThe most likely cause of the anomaly is a tank-insulation degradation which w...', 'Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...', 'Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...', 'Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...', 'Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...'] +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 1.0 +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.167 +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:17:44 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.567, Perfect scores: 1/6 +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 5/6 +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 6.17 +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:17:44 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:17:44 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.569, Max reward: 0.766 +2025-04-04 at 04:17:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:17:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Apollo 11 mission command module hot water port composition +2025-04-04 at 04:17:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module hot water port nickel content +2025-04-04 at 04:17:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What is the composition of the stainless steel used in the command module's hot water port in scuba diving? + + +2025-04-04 at 04:17:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:17:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel content in command module hot water port +2025-04-04 at 04:17:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo command module hot water port nickel content Apollo 11 +2025-04-04 at 04:17:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:17:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +command module nickel content specific alloy materials used in astronaut water systems space mission + + +2025-04-04 at 04:17:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:17:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:17:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: low nickel content command module hot water port +2025-04-04 at 04:17:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:17:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 command module water nickel content command module hot water port +2025-04-04 at 04:17:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-04 at 04:17:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What nickel alloy composition was used in the command module from NASA? + + +2025-04-04 at 04:17:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:17:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:17:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 command module hot water system nickel content +2025-04-04 at 04:17:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:17:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What type of nickel plating was used on the Apollo command module? + + +2025-04-04 at 04:17:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:17:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:17:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar command module nickel contamination +2025-04-04 at 04:17:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:17:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What type of electroless nickel plating was used on the Apollo command module? + + +2025-04-04 at 04:17:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:17:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:17:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 hot water system nickel contamination fix +2025-04-04 at 04:17:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-04 at 04:17:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What type of nickel plating used in space exploration? + + +2025-04-04 at 04:17:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:17:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:18:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 fix for nickel contamination water port +2025-04-04 at 04:18:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:18:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What type of nickel alloy was used in the Apollo's drinking water system? + + +2025-04-04 at 04:18:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-04 at 04:18:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:18:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 hot water port nickel content contamination fix +2025-04-04 at 04:18:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:18:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What specific nickel alloy was used for the potable water system of the Apollo command module? + + +2025-04-04 at 04:18:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-04 at 04:18:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:18:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module hot water nickel contamination +2025-04-04 at 04:18:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:18:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Using NASA documentation what alloy is used to prevent ice formation in the Apollo command module? + + +2025-04-04 at 04:18:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-04 at 04:18:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:18:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 hot water containment system description +2025-04-04 at 04:18:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-04 at 04:18:04 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:18:04 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:18:04 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, False, False] +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_correctness:82 - Student lengths: [236, 311, 1898, 410, 605, 1711] +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [20, 20, 20, 20, 20, 20] +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_correctness:84 - Average student length: 861.83 +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 20.00 +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_correctness:86 - Length ratio: 43.09 +2025-04-04 at 04:18:04 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-04 at 04:18:04 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.108 Âą 0.242 +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 0.50 Âą 1.12 +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [3, 0, 0, 0, 0, 0] +2025-04-04 at 04:18:04 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-04 at 04:18:04 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-04 at 04:18:04 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-04 at 04:18:04 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 3.0 +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.500 +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.500 +2025-04-04 at 04:18:04 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.200, Perfect scores: 0/6 +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 2/6 +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 2.50 +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 0/6 +2025-04-04 at 04:18:04 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:18:04 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.178, Max reward: 0.509 +2025-04-04 at 04:18:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:18:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "resolving unit for shaft axis +2025-04-04 at 04:18:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-04 at 04:18:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: type "mechanical resolver shaft axis sensor +2025-04-04 at 04:18:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-04 at 04:18:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: axis resolver +2025-04-04 at 04:18:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:18:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: "rotor resolver vs encoder" +2025-04-04 at 04:18:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-04 at 04:18:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mechanical resolver vs position sensor vs rotary encoder +2025-04-04 at 04:18:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-04 at 04:18:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: shaft position sensor types +2025-04-04 at 04:18:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-04 at 04:18:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:18:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: half-speed resolver application in mechanical systems +2025-04-04 at 04:18:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-04 at 04:18:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: body-axis alignment system +2025-04-04 at 04:18:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:18:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: terms:half speed resolver vacuum effects +2025-04-04 at 04:18:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:18:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: half-speed resolver vs full-speed resolver +2025-04-04 at 04:18:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:18:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: axes resolvers +2025-04-04 at 04:18:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-04 at 04:18:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:18:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum testing results on half-speed resolver vacuum immunity +2025-04-04 at 04:18:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:18:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: inertial measurement unit automated aircraft +2025-04-04 at 04:18:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ + +2025-04-04 at 04:18:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft half-speed resolver wiping action +2025-04-04 at 04:18:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-04 at 04:18:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum suitability of half-speed resolver +2025-04-04 at 04:18:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:18:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: star tracker half-speed resolver +2025-04-04 at 04:18:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-04 at 04:18:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:18:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: half-speed resolver optical head design +2025-04-04 at 04:18:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-04 at 04:18:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: inertial measurement unit acceleration bias +2025-04-04 at 04:18:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-04 at 04:18:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apolo ll flight profile. +2025-04-04 at 04:18:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:18:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum sealing in optical resolvers +2025-04-04 at 04:18:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:18:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: star tracker inertial measurement unit +2025-04-04 at 04:18:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:18:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:18:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 1960s lunar module lunar lander electronics +2025-04-04 at 04:18:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-04 at 04:18:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 imu bias error +2025-04-04 at 04:18:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:18:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: design considerations for vacuum sealing in optical components +2025-04-04 at 04:18:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:18:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: inertial measurement unit spacecraft navigation +2025-04-04 at 04:18:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-04 at 04:18:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:18:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module command module alignment issues +2025-04-04 at 04:18:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-04 at 04:18:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar module navigation +2025-04-04 at 04:18:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:18:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum sealed optical components for aerospace applications +2025-04-04 at 04:18:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:18:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo mission inertial navigation +2025-04-04 at 04:18:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:18:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:18:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module and command module alignment procedures +2025-04-04 at 04:18:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:18:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 crew training geology +2025-04-04 at 04:18:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:18:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA resolver approval documentation +2025-04-04 at 04:18:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:18:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo translunar injection +2025-04-04 at 04:18:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:18:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:18:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module and command module navigation and control systems +2025-04-04 at 04:18:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-04 at 04:18:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 launch and flight operations +2025-04-04 at 04:18:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:18:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft ground testing Apollo +2025-04-04 at 04:18:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:18:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: translunar injection apollo 12 +2025-04-04 at 04:18:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:18:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:18:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module separation from command module procedures +2025-04-04 at 04:18:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-04 at 04:18:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apolo 13 crew training lunar module helium pressure +2025-04-04 at 04:18:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:18:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: post-flight analysis of Apollo 13 mission +2025-04-04 at 04:18:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:18:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:18:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module separation timing and procedures +2025-04-04 at 04:18:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:18:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 helium tank pressure range test +2025-04-04 at 04:18:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:18:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission resolution +2025-04-04 at 04:18:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:18:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:18:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module and Earth re-entry procedures +2025-04-04 at 04:18:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:18:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo training for lunar missions +2025-04-04 at 04:18:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:18:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:18:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module re-entry Earth orbital insertion maneuver +2025-04-04 at 04:18:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:18:34 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:18:34 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:18:35 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, True, False, False, False] +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_correctness:82 - Student lengths: [247, 1844, 511, 1717, 2032, 1977] +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [19, 19, 19, 19, 19, 19] +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_correctness:84 - Average student length: 1388.00 +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 19.00 +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_correctness:86 - Length ratio: 73.05 +2025-04-04 at 04:18:35 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:18:35 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.192 Âą 0.301 +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 0.83 Âą 1.46 +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [1, 0, 4, 0, 0, 0] +2025-04-04 at 04:18:35 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 12: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-04 at 04:18:35 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nTime hr:min Optian code Star used Ster angle aifference, deg Gyro torquing an...', 'Result 1:\nEarth-centered inertial coordinates. Bystem. **Change in velocity showm in bo...', 'Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...'] +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 5.0 +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.833 +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:18:35 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.300, Perfect scores: 1/6 +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.67 +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:18:35 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:18:35 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.654, Max reward: 0.789 +2025-04-04 at 04:18:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:18:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: to gather more information on what happens to contaminants when they freeze. + + +2025-04-04 at 04:18:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-04 at 04:18:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How do substances with different freezing points behave when they freeze, such as water or salt? +2025-04-04 at 04:18:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:18:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "What happens to contaminants in a frozen state?" +2025-04-04 at 04:18:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:18:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to contaminants in a medium when it freezes? +2025-04-04 at 04:18:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:18:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:18:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Effects of hydrogen freezing and vapor pressure on helium transmission +2025-04-04 at 04:18:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:18:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "behavior of contaminants in the cold vacuum of space" +2025-04-04 at 04:18:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:18:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to contaminants in a spacecraft when it reaches a temperature below the contaminant's vaporization point? +2025-04-04 at 04:18:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-04 at 04:18:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:18:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Material science: effects of low concentrations of contaminants on heat exchanger performance +2025-04-04 at 04:18:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-04 at 04:18:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "contaminant-related issues in Apollo spacecraft" +2025-04-04 at 04:18:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:18:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to contaminant buildup in a water tank when it freezes? +2025-04-04 at 04:18:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:18:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:18:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Cryogenic tank heater and switch operation +2025-04-04 at 04:18:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:18:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "hydrogen gas bubbles in a sealed liquid water system" +2025-04-04 at 04:18:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:18:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to contaminants when they freeze and become frozen in heat exchanger lines? +2025-04-04 at 04:18:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:18:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:18:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Design changes to prevent hydrogen or oxygen valve closures +2025-04-04 at 04:18:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-04 at 04:18:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "thermal conductivity of hydrogen" +2025-04-04 at 04:18:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:18:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens when a low-pressurized contaminant freezes in heat exchanger lines? +2025-04-04 at 04:18:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:18:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:18:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent stage tank shutoff valve design and installation issues +2025-04-04 at 04:18:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-04 at 04:18:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "helium's thermal conductivity" +2025-04-04 at 04:18:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:18:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens after a contaminant freezes in heat exchanger lines in a spacecraft? +2025-04-04 at 04:18:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:18:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:19:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Excessive leakage through oxygen manifold valves +2025-04-04 at 04:19:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-04 at 04:19:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "physical properties of hydrogen below 0 degrees Celsius" +2025-04-04 at 04:19:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:19:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to gas bubbles in water inside a tank when it freezes and does not get used? +2025-04-04 at 04:19:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-04 at 04:19:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:19:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Error in O-ring compression in oxygen manifold valves +2025-04-04 at 04:19:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-04 at 04:19:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "behavior of hydrogen gas in low-temperature cold soak" +2025-04-04 at 04:19:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:19:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the effect on contaminants in a tank when helium is used to cool it and heaters areused to assist evaporation? +2025-04-04 at 04:19:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:19:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:19:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "vaporization and condensation of hydrogen in low-pressure environments" +2025-04-04 at 04:19:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:19:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Do frozen contaminants on the inner wall of a tank decrease pressure? +2025-04-04 at 04:19:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:19:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:19:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "in-flight testing of Apollo spacecraft propulsion systems" +2025-04-04 at 04:19:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:19:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Do frozen contaminants reduce pressure in a tank? +2025-04-04 at 04:19:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:19:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:19:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "contamination and spacecraft propulsion systems" +2025-04-04 at 04:19:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-04 at 04:19:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Do frozen contaminants reduce pressure in a tank through condensation on the inner wall? +2025-04-04 at 04:19:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:19:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:19:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Does condensation of vapors on the inner wall of a tank stabilize pressure fluctuations? +2025-04-04 at 04:19:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:19:10 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:19:10 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:19:10 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_correctness:82 - Student lengths: [745, 1618, 1587, 365, 1328, 1341] +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [92, 92, 92, 92, 92, 92] +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_correctness:84 - Average student length: 1164.00 +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 92.00 +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_correctness:86 - Length ratio: 12.65 +2025-04-04 at 04:19:10 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-04 at 04:19:10 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.079 Âą 0.177 +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 4.17 Âą 9.32 +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 25, 0, 0, 0] +2025-04-04 at 04:19:10 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-04 at 04:19:10 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-04 at 04:19:10 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...'] +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-04 at 04:19:10 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 12: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 3.0 +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.500 +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.500 +2025-04-04 at 04:19:10 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.267, Perfect scores: 1/6 +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 2/6 +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 3.00 +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:19:10 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:19:10 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.229, Max reward: 0.698 +2025-04-04 at 04:19:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:19:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 launch date +2025-04-04 at 04:19:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 moon launch date saturn ib +2025-04-04 at 04:19:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-04 at 04:19:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 mission launch date +2025-04-04 at 04:19:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:19:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 4 launch date +2025-04-04 at 04:19:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 launch date +2025-04-04 at 04:19:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:19:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: results including Apollo 4 launch information +2025-04-04 at 04:19:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 mission launch date +2025-04-04 at 04:19:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:19:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 mission training Apollo 13 mission +2025-04-04 at 04:19:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:19:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 4 mission launch date +2025-04-04 at 04:19:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:19:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 launch date +2025-04-04 at 04:19:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:19:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 launch date november 1967 +2025-04-04 at 04:19:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: йО Apollo 4 mission launch +2025-04-04 at 04:19:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:19:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 4 mission training Apollo 13 +2025-04-04 at 04:19:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:19:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 4 launch date November 9 1967 +2025-04-04 at 04:19:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 lifecycle mission date +2025-04-04 at 04:19:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:19:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: when did the uncrewed apollo 4 mission take place +2025-04-04 at 04:19:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:19:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 mission training Apollo 3 mission +2025-04-04 at 04:19:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:19:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: confirm Apollo 4 launch date is April 11 1970 +2025-04-04 at 04:19:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 launch mission date 1967 +2025-04-04 at 04:19:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:19:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: when was the uncrewed apollo 4 mission +2025-04-04 at 04:19:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:19:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo training April 1969 +2025-04-04 at 04:19:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 4 mission November 9 1967 +2025-04-04 at 04:19:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no change +2025-04-04 at 04:19:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:19:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:19:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: when was the apollo 4 launch +2025-04-04 at 04:19:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 training JFR +2025-04-04 at 04:19:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 4 mission launch date +2025-04-04 at 04:19:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:19:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 launch date +2025-04-04 at 04:19:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:19:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 training and saturn SA-506 +2025-04-04 at 04:19:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:19:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V uncrewed test launch date +2025-04-04 at 04:19:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ + +2025-04-04 at 04:19:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 10 and 11 launch dates +2025-04-04 at 04:19:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:19:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:19:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 training +2025-04-04 at 04:19:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:19:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 launch date +2025-04-04 at 04:19:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:19:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 Saturn III training +2025-04-04 at 04:19:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:19:31 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:19:31 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:19:31 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, False, False, True, False] +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_correctness:82 - Student lengths: [223, 218, 178, 2032, 285, 2017] +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_correctness:84 - Average student length: 825.50 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 12.00 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_correctness:86 - Length ratio: 68.79 +2025-04-04 at 04:19:31 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.833, Valid formats: 5.0/6 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_format:228 - Responses ending properly: 6/6 +2025-04-04 at 04:19:31 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.425 Âą 0.437 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.67 Âą 2.92 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 3, 6, 0, 7, 0] +2025-04-04 at 04:19:31 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-04 at 04:19:31 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 5.0 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.833 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:19:31 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.767, Perfect scores: 4/6 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 5/6 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 4.67 +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 4/6 +2025-04-04 at 04:19:31 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:19:31 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.217, Max reward: 0.563 +2025-04-04 at 04:19:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:19:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how long has planetary photography been a well-established field of study or practice +2025-04-04 at 04:19:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-04 at 04:19:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "history of weather photography duration" +2025-04-04 at 04:19:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Marshall Space Flight Center, Kennedy Space Center, Manned Spacecraft Center: Analysis of Apollo l2 Lightning Incident, MSC-01540. February 1970. + +ICSU/IUGG Committee on Atmospheric Sciences: Report of the Study Conference on the Global Atmospheric Research Program, 1967. + +Bulletin of the American Meteorological Society, Vol. 50, No. 7: Cloud Height Contouring from Apollo 6 Photography, by V. S. Whitehead, I. D. Browne, and J. G. Garcia. 1969. + +Defense Supply Agency, Washington, D. C.: Military Standardization Handbook_ Optical Design, MIL HDBK-14l. 1962. + +NASA Headquarters: Apollo Flight Mission Assignments. OMSF M-D MA500-11 (SE 010-000-1). 0ctober 1969. + +Manned Spacecraft Center: Mission Requirement, H-2 Type Mission (Lunar Landing). SPD9-R-053. November 10, 1969. + +APOLLO SPACECRAFT FLIGHT HISTORY + +(Continued from inside front cover) +------ + +2025-04-04 at 04:19:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Earth observation program origin and duration +2025-04-04 at 04:19:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-04 at 04:19:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how long did earth have a photosphere +2025-04-04 at 04:19:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-04 at 04:19:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:19:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how long was earth weather photography conducted after translunar injection during a space mission +2025-04-04 at 04:19:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-04 at 04:19:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "history of Earth's weather photography" +2025-04-04 at 04:19:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-04 at 04:19:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 23 6 hour navigation +2025-04-04 at 04:19:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:19:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo mission earth shadow duration" +2025-04-04 at 04:19:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:19:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo navigation maneuver +2025-04-04 at 04:19:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:19:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:19:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 descent phase photography duration" +2025-04-04 at 04:19:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:19:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:19:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 lunar surface photography duration" +2025-04-04 at 04:19:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:19:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:19:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 lunar surface photography duration" +2025-04-04 at 04:19:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:19:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:19:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 lunar surface lunar surface photography duration" +2025-04-04 at 04:19:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:19:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:19:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "duration of moon photography missions" +2025-04-04 at 04:19:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-04 at 04:19:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:19:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo mission duration" +2025-04-04 at 04:19:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:19:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:19:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 mission duration" +2025-04-04 at 04:19:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-04 at 04:19:50 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:19:50 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:19:50 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, True, False, False, False] +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_correctness:82 - Student lengths: [308, 1726, 746, 1211, 279, 246] +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [21, 21, 21, 21, 21, 21] +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_correctness:84 - Average student length: 752.67 +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 21.00 +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_correctness:86 - Length ratio: 35.84 +2025-04-04 at 04:19:50 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-04 at 04:19:50 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.142 Âą 0.205 +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 0.50 Âą 0.76 +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [2, 0, 1, 0, 0, 0] +2025-04-04 at 04:19:50 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Marshall Space Flight Center, Kennedy Space Center, Manned Spacecraft Center: Analysis of Apollo l2 Lightning Incident, MSC-01540. February 1970. + +ICSU/IUGG Committee on Atmospheric Sciences: Report of the Study Conference on the Global Atmospheric Research Program, 1967. + +Bulletin of the American Meteorological Society, Vol. 50, No. 7: Cloud Height Contouring from Apollo 6 Photography, by V. S. Whitehead, I. D. Browne, and J. G. Garcia. 1969. + +Defense Supply Agency, Washington, D. C.: Military Standardization Handbook_ Optical Design, MIL HDBK-14l. 1962. + +NASA Headquarters: Apollo Flight Mission Assignments. OMSF M-D MA500-11 (SE 010-000-1). 0ctober 1969. + +Manned Spacecraft Center: Mission Requirement, H-2 Type Mission (Lunar Landing). SPD9-R-053. November 10, 1969. + +APOLLO SPACECRAFT FLIGHT HISTORY + +(Continued from inside front cover) +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-04 at 04:19:50 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-04 at 04:19:50 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:19:50 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.333, Perfect scores: 0/6 +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.67 +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:19:50 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:19:50 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.264, Max reward: 0.729 +2025-04-04 at 04:19:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:19:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Johnson Space Center vs Kennedy Space Center +2025-04-04 at 04:19:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:19:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: johnson space center location +2025-04-04 at 04:19:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-04 at 04:19:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center location +2025-04-04 at 04:19:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-04 at 04:19:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Location of NASA Johnson Space Center +2025-04-04 at 04:19:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:19:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center location +2025-04-04 at 04:19:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-04 at 04:19:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center location +2025-04-04 at 04:19:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-04 at 04:19:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:19:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center location historical documents +2025-04-04 at 04:19:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:19:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center vs Lyndon B. Johnson Space Center +2025-04-04 at 04:19:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:19:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: US Government websites nasa.msc.nasa.gov +2025-04-04 at 04:19:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:19:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Houston Manned Spacecraft Center mailing address +2025-04-04 at 04:19:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:19:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Johnson Space Center Houston address +2025-04-04 at 04:19:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:19:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:19:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center location +2025-04-04 at 04:19:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-04 at 04:19:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Johnson Space Center +2025-04-04 at 04:19:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:19:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Johnson Space Center 77058 +2025-04-04 at 04:19:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:19:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:20:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lyndon B. Johnson Space Center address +2025-04-04 at 04:20:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:20:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:20:03 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:20:03 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:20:03 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, True, False, True, True] +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.83 +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_correctness:82 - Student lengths: [287, 405, 221, 255, 151, 402] +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [14, 14, 14, 14, 14, 14] +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_correctness:84 - Average student length: 286.83 +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 14.00 +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_correctness:86 - Length ratio: 20.49 +2025-04-04 at 04:20:03 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:20:03 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.421 Âą 0.254 +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.50 Âą 1.50 +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 2/6 +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [4, 1, 3, 0, 3, 4] +2025-04-04 at 04:20:03 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-04 at 04:20:03 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...'] +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:20:03 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nPOSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION\n\nNASA-Mann...', 'Result 1:\nPOSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION\n\nNASA-Mann...'] +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:20:03 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...', 'Result 1:\nPOSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION\n\nNASA-Mann...'] +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 3.0 +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.500 +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.500 +2025-04-04 at 04:20:03 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.600, Perfect scores: 2/6 +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 5/6 +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.33 +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 3/6 +2025-04-04 at 04:20:03 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:20:03 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.452, Max reward: 0.696 +2025-04-04 at 04:20:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:20:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the time of S-II crossbeam oscillations peak amplitude +33.7g? +2025-04-04 at 04:20:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "S-II crossbeam oscillations +33.7g peak amplitude" +2025-04-04 at 04:20:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: when did s ii crossbeam oscillations peak amplitude +33.7g +2025-04-04 at 04:20:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-II crossbeam oscillations peak amplitude 33.7g +2025-04-04 at 04:20:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Space Shuttle S-2 orbital re-entry crossbeam oscillations amplitude +2025-04-04 at 04:20:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:20:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:20:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Verify Apollo 12 S-II oscillation event timing 330.6 seconds +2025-04-04 at 04:20:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-04 at 04:20:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 S-II crossbeam oscillations 33.7g peak time" +2025-04-04 at 04:20:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-04 at 04:20:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what time did the S-II crossbeam oscillations peak at +33.7g apollo 13 +2025-04-04 at 04:20:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-04 at 04:20:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-II crossbeam oscillations 33.7g +2025-04-04 at 04:20:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-04 at 04:20:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-II crossbeam oscillations time = 330.6 seconds +2025-04-04 at 04:20:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:20:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Verify seismic data after S-IVB impact timing +2025-04-04 at 04:20:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:20:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the meaning of 330.6 seconds in a 13-minute clock time +2025-04-04 at 04:20:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +-1.2 40.4 +0.4 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:20:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-II engine 5 cutoff time +2025-04-04 at 04:20:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:20:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 00:05:32 S-II premature engine shutdown occur +2025-04-04 at 04:20:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-04 at 04:20:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:20:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Verify connection between S-IVB impact and S-II oscillation timing +2025-04-04 at 04:20:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-04 at 04:20:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 13 minutes past 0 hours, 0 minutes, and 0 seconds in MET apollo 13 +2025-04-04 at 04:20:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:20:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-II engine 5 cutoff time +2025-04-04 at 04:20:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:20:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-II premature shutdown oxygen tank 1 and 2 +2025-04-04 at 04:20:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-04 at 04:20:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:20:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-II oscillation time before S-IVB staging +2025-04-04 at 04:20:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-04 at 04:20:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what time is 13 minutes past 19:00 on April 11, 1970 +2025-04-04 at 04:20:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:20:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-II crossbeam oscillations oxygen tank 2 short +2025-04-04 at 04:20:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ + +2025-04-04 at 04:20:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:20:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-II crossbeam oscillations time 330.6 seconds +2025-04-04 at 04:20:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:20:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-II peak amplitude 33.7g time +2025-04-04 at 04:20:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-04 at 04:20:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:20:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-II 00:05:32 +2025-04-04 at 04:20:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +14.2.2。 +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:20:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:20:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-II 00:12:30 +2025-04-04 at 04:20:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-04 at 04:20:26 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:20:26 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:20:26 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, False, False, False, False] +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_correctness:82 - Student lengths: [568, 274, 152, 108, 1227, 384] +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_correctness:84 - Average student length: 452.17 +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 13.00 +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_correctness:86 - Length ratio: 34.78 +2025-04-04 at 04:20:26 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-04 at 04:20:26 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.375 Âą 0.398 +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 1.83 Âą 2.03 +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [5, 2, 0, 4, 0, 0] +2025-04-04 at 04:20:26 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +-1.2 40.4 +0.4 +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +a。 Helium l on quads B and D b。 Helium 2 on quad D C. Secondary propellant valves on quads A and C. + +Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power. + +The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred. +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +14.2.2。 +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-04 at 04:20:26 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 5.0 +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.833 +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:20:26 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.433, Perfect scores: 0/6 +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 2.67 +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:20:26 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:20:26 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.516, Max reward: 0.818 +2025-04-04 at 04:20:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:20:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: adverse effects of extraneous firings in turbine applications +2025-04-04 at 04:20:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-04 at 04:20:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what are the adverse effects of two extraneous firings +2025-04-04 at 04:20:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-04 at 04:20:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: extraneous firings figure 5.6-1 adverse effects +2025-04-04 at 04:20:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:20:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (extraneous firings figure 5.6-1 effects) +2025-04-04 at 04:20:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:20:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"adverse effects of electrical sparks in circuit 5.6-1 figure" +2025-04-04 at 04:20:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-04 at 04:20:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:20:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: extraneous firings apollo accident report +2025-04-04 at 04:20:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: adverse effects of extraneous firings on Apollo spacecraft +2025-04-04 at 04:20:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:20:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command and service module separation adverse effects +2025-04-04 at 04:20:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-04 at 04:20:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: extraneous firings in figure 5.6-1 adverse effects +2025-04-04 at 04:20:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-04 at 04:20:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"adverse effects of burning insulation in spacecraft electrical conduit" +2025-04-04 at 04:20:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-04 at 04:20:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:20:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 extraneous firing 5 +2025-04-04 at 04:20:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:20:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: adverse effects of high vibration on Apollo spacecraft +2025-04-04 at 04:20:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-04 at 04:20:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command and service module postrecovery inspection discrepancies +2025-04-04 at 04:20:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-04 at 04:20:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gyrocompassing misalignment extraneous firings adverse effects +2025-04-04 at 04:20:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-04 at 04:20:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"mechanism of battery degradation during cold soak" +2025-04-04 at 04:20:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Evidence indicates that battery 2 may have experienced an electrical fault of some type. The most probable condition is electrolyte leaking from one or more cells and bridging the high-voltage or low-voltage terminal to the battery case (fig. 14-17). This bridging results in water electrolysis and subsequent ignition of the hydrogen and oxygen so generated. The accompanying "explosion" would then blow off or rupture the seal of the battery lid and cause both a thump and venting of the free liquids in the battery case, resulting in "snowflakes." + +Postflight tests have shown the following: +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-04 at 04:20:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:20:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 electrical charge separation launch hazard +2025-04-04 at 04:20:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:20:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effects of launch oscillations on Apollo spacecraft structural components +2025-04-04 at 04:20:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ + +2025-04-04 at 04:20:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command and service module structural mechanical systems anomalies +2025-04-04 at 04:20:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +The crew reported a pitch-up disturbance torque was exerted on the command module soon after undocking until the beginning of entry. Most of this time, only low-bit-rate telemetry was available and therefore a detailed analysis is impossible. A 20-minute segment of high-bit-rate data was received just prior to entry, and an unaccountable pitch-up torque of 0.00l deg/sec2 was observed. The possible contributing causes for this torque could have been gravity gradients, atmospheric trimming, venting through the umbilical, venting through the tunnel hatch, and a gradual propellant leak. However, none of these is considered to have been a single cause, and either a combination of these causes was present Or some undetermined venting took place. +------ + +2025-04-04 at 04:20:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"consequences of battery current surge from electrical fault" +2025-04-04 at 04:20:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +A battery overcurrent can be ruled out because automatic removal of the battery from the bus would have occurred. + +A reverse-current condition can be ruled out because, if the battery is removed from and reapplied to the bus, the reverse-current circuit has a built-in delay of about 5 seconds before the reverse-current relay is again activated to illuminate the light. Battery power was removed from and replaced on the bus in flight, and the light immediately illuminated again when the battery was reconnected. + +An over-temperature condition can be ruled out because, after the battery was replaced on the bus, the light remained illuminated for a brief period and then began flickering intermittently. A flickering light cannot be caused by the temperature sensing switch because of a temperature hysteresis of approximately $_{20}\circ$ F in the switch. The water glycol loop temperature also indicated that the battery temperature was normal. +------ + +2025-04-04 at 04:20:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:20:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Extraneous firing Apollo 13 launch operation +2025-04-04 at 04:20:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:20:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: low-frequency vibrations and S-II engine failure in Apollo spacecraft +2025-04-04 at 04:20:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch-up torque command module entry phase stability +2025-04-04 at 04:20:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew reported a pitch-up disturbance torque was exerted on the command module soon after undocking until the beginning of entry. Most of this time, only low-bit-rate telemetry was available and therefore a detailed analysis is impossible. A 20-minute segment of high-bit-rate data was received just prior to entry, and an unaccountable pitch-up torque of 0.00l deg/sec2 was observed. The possible contributing causes for this torque could have been gravity gradients, atmospheric trimming, venting through the umbilical, venting through the tunnel hatch, and a gradual propellant leak. However, none of these is considered to have been a single cause, and either a combination of these causes was present Or some undetermined venting took place. +------ +Result 2: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ + +2025-04-04 at 04:20:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"challenger electrical fault battery 1" +2025-04-04 at 04:20:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ + +2025-04-04 at 04:20:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:20:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: extraneous electrical discharges space missions +2025-04-04 at 04:20:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-04 at 04:20:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: high-amplitude vibration and engine shutdown during S-II thrust phase Apollo 13 +2025-04-04 at 04:20:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gimbal motor voltage regulator differences command and service module +2025-04-04 at 04:20:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-04 at 04:20:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +" corrective action for Apollo 13 battery malfunction" +2025-04-04 at 04:20:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:20:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:20:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 extraneous firing effects +2025-04-04 at 04:20:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:20:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: chapter 13 adverse effects Apollo 13 +2025-04-04 at 04:20:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-04 at 04:20:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gimbal motor control and battery voltage regulation anomalies effect on structural mechanics +2025-04-04 at 04:20:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The recurrence of the problem under almost identical circumstances during Apollo l3 indicates that the cause is more likely generic than random and that it is time or vacuum dependent. The susceptibility of the shaft axis rather than the trunnion axis also tends to absolve components common to both axes, such as. the electronics and the motor drive amplifier. The shaft loop has been shown to be more sensitive than the trunnion to harmonics of the 800-hertz reference voltages introduced into the forward loop; however, because the level of the required null offset voltage is well above that available by induction, this mechanism is considered unlikely. +------ + +2025-04-04 at 04:20:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"apod 13 crew survival strategy" +2025-04-04 at 04:20:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-04 at 04:20:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:20:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electrical charges launch conditions +2025-04-04 at 04:20:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-04 at 04:20:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 effects of extraneous firings on crew physiological data +2025-04-04 at 04:20:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-04 at 04:20:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electronic component issues motor drive amplifier structural mechanics +2025-04-04 at 04:20:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The recurrence of the problem under almost identical circumstances during Apollo l3 indicates that the cause is more likely generic than random and that it is time or vacuum dependent. The susceptibility of the shaft axis rather than the trunnion axis also tends to absolve components common to both axes, such as. the electronics and the motor drive amplifier. The shaft loop has been shown to be more sensitive than the trunnion to harmonics of the 800-hertz reference voltages introduced into the forward loop; however, because the level of the required null offset voltage is well above that available by induction, this mechanism is considered unlikely. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-04 at 04:20:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"apod 13 primary recovery operations" +2025-04-04 at 04:20:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-04 at 04:20:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:20:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electrostatic charges spacecraft operations +2025-04-04 at 04:20:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-04 at 04:20:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: zero optics mode circuitry residue impact on motor drive amplifier +2025-04-04 at 04:20:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-04 at 04:20:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"effects of apod 13 limited Saturn l receptor filter failure " +2025-04-04 at 04:20:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-04 at 04:20:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:20:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: zero optics mode circuitry contamination investigation +2025-04-04 at 04:20:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:20:56 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:20:56 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:20:56 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, False, True, False, False] +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1893, 663, 1423, 506, 1660, 849] +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [41, 41, 41, 41, 41, 41] +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_correctness:84 - Average student length: 1165.67 +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 41.00 +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_correctness:86 - Length ratio: 28.43 +2025-04-04 at 04:20:56 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_format:228 - Responses ending properly: 6/6 +2025-04-04 at 04:20:56 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.267 Âą 0.387 +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 1.83 Âą 2.97 +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 8, 0, 3, 0, 0] +2025-04-04 at 04:20:56 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +The crew reported a pitch-up disturbance torque was exerted on the command module soon after undocking until the beginning of entry. Most of this time, only low-bit-rate telemetry was available and therefore a detailed analysis is impossible. A 20-minute segment of high-bit-rate data was received just prior to entry, and an unaccountable pitch-up torque of 0.00l deg/sec2 was observed. The possible contributing causes for this torque could have been gravity gradients, atmospheric trimming, venting through the umbilical, venting through the tunnel hatch, and a gradual propellant leak. However, none of these is considered to have been a single cause, and either a combination of these causes was present Or some undetermined venting took place. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The crew reported a pitch-up disturbance torque was exerted on the command module soon after undocking until the beginning of entry. Most of this time, only low-bit-rate telemetry was available and therefore a detailed analysis is impossible. A 20-minute segment of high-bit-rate data was received just prior to entry, and an unaccountable pitch-up torque of 0.00l deg/sec2 was observed. The possible contributing causes for this torque could have been gravity gradients, atmospheric trimming, venting through the umbilical, venting through the tunnel hatch, and a gradual propellant leak. However, none of these is considered to have been a single cause, and either a combination of these causes was present Or some undetermined venting took place. +------ +Result 2: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ +Result 2: +The recurrence of the problem under almost identical circumstances during Apollo l3 indicates that the cause is more likely generic than random and that it is time or vacuum dependent. The susceptibility of the shaft axis rather than the trunnion axis also tends to absolve components common to both axes, such as. the electronics and the motor drive amplifier. The shaft loop has been shown to be more sensitive than the trunnion to harmonics of the 800-hertz reference voltages introduced into the forward loop; however, because the level of the required null offset voltage is well above that available by induction, this mechanism is considered unlikely. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The recurrence of the problem under almost identical circumstances during Apollo l3 indicates that the cause is more likely generic than random and that it is time or vacuum dependent. The susceptibility of the shaft axis rather than the trunnion axis also tends to absolve components common to both axes, such as. the electronics and the motor drive amplifier. The shaft loop has been shown to be more sensitive than the trunnion to harmonics of the 800-hertz reference voltages introduced into the forward loop; however, because the level of the required null offset voltage is well above that available by induction, this mechanism is considered unlikely. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Evidence indicates that battery 2 may have experienced an electrical fault of some type. The most probable condition is electrolyte leaking from one or more cells and bridging the high-voltage or low-voltage terminal to the battery case (fig. 14-17). This bridging results in water electrolysis and subsequent ignition of the hydrogen and oxygen so generated. The accompanying "explosion" would then blow off or rupture the seal of the battery lid and cause both a thump and venting of the free liquids in the battery case, resulting in "snowflakes." + +Postflight tests have shown the following: +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +A battery overcurrent can be ruled out because automatic removal of the battery from the bus would have occurred. + +A reverse-current condition can be ruled out because, if the battery is removed from and reapplied to the bus, the reverse-current circuit has a built-in delay of about 5 seconds before the reverse-current relay is again activated to illuminate the light. Battery power was removed from and replaced on the bus in flight, and the light immediately illuminated again when the battery was reconnected. + +An over-temperature condition can be ruled out because, after the battery was replaced on the bus, the light remained illuminated for a brief period and then began flickering intermittently. A flickering light cannot be caused by the temperature sensing switch because of a temperature hysteresis of approximately $_{20}\circ$ F in the switch. The water glycol loop temperature also indicated that the battery temperature was normal. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted. + +This anomaly is closed. + +14.2.3 Descent Battery 2 Malfunction Light On + +The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8. + +NASA-S-70-5860 + + + +Figure l4-l8.- Battery 2 malfunction circuit. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-04 at 04:20:56 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe next series of events occurred within a fraction of a second between the ...', 'Result 1:\nThe next series of events occurred within a fraction of a second between the ...', 'Result 1:\nEvidence indicates that battery 2 may have experienced an electrical fault of...', 'Result 1:\nThe data show that descent battery 2 experienced at least a 60-ampere current...', 'Result 1:\nThe data show that descent battery 2 experienced at least a 60-ampere current...', 'Result 1:\nThe designs of other Apollo batteries have been reevaluated, and all are cons...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nThe Mission Control Center and the Manned Space Flight Network provided excel...', 'Result 1:\nAn investigation conducted after Apollo l2 did not identify a definite source...'] +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-04 at 04:20:56 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:20:56 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.267, Perfect scores: 0/6 +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 2.83 +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 0/6 +2025-04-04 at 04:20:56 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:20:56 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.442, Max reward: 0.813 +2025-04-04 at 04:20:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:21:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb engine cutoff time +2025-04-04 at 04:21:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB engine cutoff time +2025-04-04 at 04:21:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: when did S-IVB engine cutoff occur during the Apollo mission +2025-04-04 at 04:21:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the S-IVB engine cutoff time for the Apollo missions +2025-04-04 at 04:21:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the cutoff time for the S-IVB engine on the specific NASA mission where 'S-IVB' is used? +2025-04-04 at 04:21:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: search "S-IVB engine cutoff time" +2025-04-04 at 04:21:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:21:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb engine cutoff apollo 12 +2025-04-04 at 04:21:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:03 | INFO | src.agent:run_tool_calls:115 - īŋŊīŋŊīŋŊ Search Query: Apollo 12 S-IVB engine cutoff +2025-04-04 at 04:21:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar impact time S-IVB engine cutoff +2025-04-04 at 04:21:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:21:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the mission that corresponds to the S-IVB impact described in section ll.0? +2025-04-04 at 04:21:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:21:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB engine cutoff time Apollo +2025-04-04 at 04:21:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:21:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb engine cutoff apollo 12 +2025-04-04 at 04:21:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-IVB engine burn +2025-04-04 at 04:21:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the time of S-IVB engine cutoff in 24-hour format +2025-04-04 at 04:21:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the launch date of the spacecraft associated with the S-IVB engine cutoff at 00:12:30? +2025-04-04 at 04:21:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-04 at 04:21:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB engine cutoff time Apollo +2025-04-04 at 04:21:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:21:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ariane iv s ivb engine cutoff time +2025-04-04 at 04:21:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 engine shutdown +2025-04-04 at 04:21:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ + +2025-04-04 at 04:21:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: when did S-IVB engine cutoff occur +2025-04-04 at 04:21:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:21:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the S-IVB engine cutoff event that occurred before the initial S-IVB engine ignition listed at 00:12:30 in the launch sequence? +2025-04-04 at 04:21:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-04 at 04:21:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:21:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb engine cutoff time apollo 12 +2025-04-04 at 04:21:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-IVB engine burn duration +2025-04-04 at 04:21:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb cutoff times +2025-04-04 at 04:21:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-04 at 04:21:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the sequence of S-IC engine cutoff with corresponding other events leading up to S-II engine ignition on the Apollo 11 mission? +2025-04-04 at 04:21:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:21:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ariane iv s iv s ivb engine cutoff +2025-04-04 at 04:21:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-IVB ignition time +2025-04-04 at 04:21:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb engine cutoff specific time +2025-04-04 at 04:21:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-04 at 04:21:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the S-IVB ignition time for the Apollo 11 mission? +2025-04-04 at 04:21:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:21:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ariane iv engine cutoff times +2025-04-04 at 04:21:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:21:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-IVB cutoff time +2025-04-04 at 04:21:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:21:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb engine cutoff time +2025-04-04 at 04:21:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the sequence of events following S-IVB ignition that could include the S-IVB cutoff for the Apollo 11 mission? +2025-04-04 at 04:21:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:21:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ariane iv l2 s ivb engine ignition and shutdown +2025-04-04 at 04:21:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the first event after S-IVB ignition that could potentially include the S-IVB engine cutoff for the Apollo 11 mission? +2025-04-04 at 04:21:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:21:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ariane iv engine 5 shutdown s ivb +2025-04-04 at 04:21:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the S-IVB separation events for the Apollo 11 mission? +2025-04-04 at 04:21:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ + +2025-04-04 at 04:21:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:21:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ariane iv s ivb engine start s ivb cut off +2025-04-04 at 04:21:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:21:22 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:21:22 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:21:22 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, False, True, False, True] +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1368, 372, 112, 134, 1467, 162] +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_correctness:84 - Average student length: 602.50 +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 8.00 +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_correctness:86 - Length ratio: 75.31 +2025-04-04 at 04:21:22 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_format:228 - Responses ending properly: 6/6 +2025-04-04 at 04:21:22 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.325 Âą 0.368 +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 1.83 Âą 2.54 +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 7, 0, 1, 0, 3] +2025-04-04 at 04:21:22 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 6.0 +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 1.000 +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.000 +2025-04-04 at 04:21:22 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.500, Perfect scores: 2/6 +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 5/6 +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 2.83 +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:21:22 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:21:22 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.487, Max reward: 0.844 +2025-04-04 at 04:21:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:21:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How many mission reports published Apollo 8 +2025-04-04 at 04:21:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:21:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 mission mission reports +2025-04-04 at 04:21:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission historic reports +2025-04-04 at 04:21:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission reports +2025-04-04 at 04:21:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:21:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Briefing on number of Apollo 8 mission reports +2025-04-04 at 04:21:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:21:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 mission reports NASA +2025-04-04 at 04:21:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:21:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 mission mission report tally +2025-04-04 at 04:21:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission technical reports +2025-04-04 at 04:21:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:21:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How many mission reports available Apollo 8 +2025-04-04 at 04:21:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:21:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 mission reports count +2025-04-04 at 04:21:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: count of evaluations for apollo 8 +2025-04-04 at 04:21:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission technical reports sorted +2025-04-04 at 04:21:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:21:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:21:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Report count by organization or author on Apollo 8 mission +2025-04-04 at 04:21:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:21:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 mission reports archives +2025-04-04 at 04:21:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 numerical documentation +2025-04-04 at 04:21:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 transearth injection navigation +2025-04-04 at 04:21:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:21:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:21:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 mission reports total +2025-04-04 at 04:21:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 final technical analysis +2025-04-04 at 04:21:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 transearth injection procedure +2025-04-04 at 04:21:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:21:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:21:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 mission reports official nasa +2025-04-04 at 04:21:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:21:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: supplementary documentation for Apollo 8 +2025-04-04 at 04:21:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 transearth injection actual mission details +2025-04-04 at 04:21:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-04 at 04:21:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:21:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 mission reports nasa documentation +2025-04-04 at 04:21:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 incident report +2025-04-04 at 04:21:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:21:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 transearth injection +2025-04-04 at 04:21:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-04 at 04:21:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:21:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apore 8 mission reports archives total +2025-04-04 at 04:21:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:21:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: post-mission analysis for apollo 8 +2025-04-04 at 04:21:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:21:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 transearth injection commander +2025-04-04 at 04:21:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:21:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:21:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 supplemental reports table E-I +2025-04-04 at 04:21:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:21:46 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:21:46 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:21:46 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_correctness:82 - Student lengths: [28, 233, 1339, 421, 2033, 1762] +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [1, 1, 1, 1, 1, 1] +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_correctness:84 - Average student length: 969.33 +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 1.00 +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_correctness:86 - Length ratio: 969.33 +2025-04-04 at 04:21:46 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-04 at 04:21:46 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.000 Âą 0.000 +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 0.00 Âą 0.00 +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 0, 0, 0, 0] +2025-04-04 at 04:21:46 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-04 at 04:21:46 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-04 at 04:21:46 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:21:46 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.200, Perfect scores: 0/6 +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 2/6 +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.83 +2025-04-04 at 04:21:46 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 0/6 +2025-04-04 at 04:21:46 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:21:47 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.095, Max reward: 0.506 +2025-04-04 at 04:21:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:21:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: when was the AS-201 SC-009 Supercircular entry located February 26 1966 +2025-04-04 at 04:21:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-04 at 04:21:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aslam sc 009 supercircular entry february 26 1966 +2025-04-04 at 04:21:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-04 at 04:21:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:21:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of AS-201 SC-009 Supercircular entry February 26 1966 +2025-04-04 at 04:21:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-04 at 04:21:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of supercircular entry 10018 +2025-04-04 at 04:21:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ + +2025-04-04 at 04:21:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:21:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 Supercircular entry launch site +2025-04-04 at 04:21:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ + +2025-04-04 at 04:21:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: northern mexico location white sands missile range +2025-04-04 at 04:21:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-04 at 04:21:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:21:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 Supercircular entry location confirmed +2025-04-04 at 04:21:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ + +2025-04-04 at 04:21:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:21:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 entry Cape Kennedy +2025-04-04 at 04:21:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-04 at 04:21:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:22:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 entry +2025-04-04 at 04:22:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-04 at 04:22:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:22:01 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:22:01 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:22:01 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, True, False, False, False] +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_correctness:82 - Student lengths: [250, 520, 320, 427, 628, 367] +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [7, 7, 7, 7, 7, 7] +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_correctness:84 - Average student length: 418.67 +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 7.00 +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_correctness:86 - Length ratio: 59.81 +2025-04-04 at 04:22:01 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-04 at 04:22:01 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.267 Âą 0.387 +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 1.50 Âą 2.29 +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [6, 0, 3, 0, 0, 0] +2025-04-04 at 04:22:01 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ + +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ + +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-04 at 04:22:01 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ + +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-04 at 04:22:01 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-04 at 04:22:01 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-04 at 04:22:01 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 2.0 +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.333 +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:22:01 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.333, Perfect scores: 2/6 +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 2/6 +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 1.17 +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:22:01 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:22:01 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.234, Max reward: 0.741 +2025-04-04 at 04:22:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:22:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space mission S-IVB vibration duration +2025-04-04 at 04:22:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the duration of a specific vibration experienced during the S-IVB mission? +2025-04-04 at 04:22:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space vibration duration S-IVB launch +2025-04-04 at 04:22:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:22:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollomissions s ivb vibration duration +2025-04-04 at 04:22:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: the duration of S-IVB vibrations on Viking 1 and Viking 2 missions +2025-04-04 at 04:22:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB vibration duration +2025-04-04 at 04:22:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:22:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:22:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb vibration apollo 13 +2025-04-04 at 04:22:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:22:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the duration of vibration on S-IVB during the Apollo 8 mission +2025-04-04 at 04:22:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB launch vibration +2025-04-04 at 04:22:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:22:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollomissions s ivb vibration duration +2025-04-04 at 04:22:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: range of S-IVB vibration during launch phase +2025-04-04 at 04:22:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:22:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB vibration duration range lunar mission +2025-04-04 at 04:22:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-04 at 04:22:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:22:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb vibration apollo 13 orbital phase +2025-04-04 at 04:22:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:22:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration event during Apollo 13 +2025-04-04 at 04:22:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:22:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB transgression time 69-122 and 122-143 +2025-04-04 at 04:22:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollomissions s ivb vibration duration after s ii stage +2025-04-04 at 04:22:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:22:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: frequency of S-IVB vibration +2025-04-04 at 04:22:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:22:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration duration +2025-04-04 at 04:22:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:22:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:22:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration event during Apollo 13 mission's launch or first 143 hours +2025-04-04 at 04:22:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:22:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-II S-IVB vibration after ignition +2025-04-04 at 04:22:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:22:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (69-122 hours and 122-143 hours) S-IVB vibration +2025-04-04 at 04:22:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration duration S-IVB +2025-04-04 at 04:22:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:22:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:22:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 S-IVB vibration 69-143 hours +2025-04-04 at 04:22:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:22:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-II S-IVB vibration after 00:05:32 +2025-04-04 at 04:22:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:22:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration - longitudinal oscillations at S-IC staging +2025-04-04 at 04:22:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration duration normal +2025-04-04 at 04:22:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:22:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:22:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration at Apollo 13 122 to 143 hours +2025-04-04 at 04:22:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:22:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration duration after 00:05:32 +2025-04-04 at 04:22:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:22:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration - longitudinal oscillations at S-II staging - engine 5 shutdown +2025-04-04 at 04:22:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:22:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration duration 122 hours +2025-04-04 at 04:22:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:22:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration characteristics Apollo 13 122 to 143 hours +2025-04-04 at 04:22:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:22:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration duration 25 minutes +2025-04-04 at 04:22:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:22:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration - S-II engine 5 shutdown at 09:05:32 +2025-04-04 at 04:22:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:22:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB engine vibration Apollo 13 122 to 143 hours +2025-04-04 at 04:22:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB vibration period 69 to 143 +2025-04-04 at 04:22:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:22:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration - Apollo 8 +2025-04-04 at 04:22:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:22:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 engine vibration discrepancy +2025-04-04 at 04:22:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ +Result 2: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ + +2025-04-04 at 04:22:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB vibration period 69 to 143 hours +2025-04-04 at 04:22:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:22:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: high-frequency S-IVB vibration - 122-143 hours +2025-04-04 at 04:22:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:24 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:22:24 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:22:24 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, True, False, True] +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_correctness:82 - Student lengths: [657, 1838, 1867, 344, 2045, 266] +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [15, 15, 15, 15, 15, 15] +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_correctness:84 - Average student length: 1169.50 +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 15.00 +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_correctness:86 - Length ratio: 77.97 +2025-04-04 at 04:22:24 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.833, Valid formats: 5.0/6 +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_format:228 - Responses ending properly: 6/6 +2025-04-04 at 04:22:24 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.375 Âą 0.388 +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.00 Âą 2.24 +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [3, 0, 0, 3, 0, 6] +2025-04-04 at 04:22:24 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ +Result 2: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 6.0 +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 1.000 +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.000 +2025-04-04 at 04:22:24 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.633, Perfect scores: 2/6 +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 5/6 +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 3.17 +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:22:24 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:22:24 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.514, Max reward: 0.830 +2025-04-04 at 04:22:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:22:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "helium tank insulation material" +2025-04-04 at 04:22:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:22:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What insulation material is commonly used in high pressure metal vessels such as helium tanks? +2025-04-04 at 04:22:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:22:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material was used for the insulation between the shells of a helium tank? +2025-04-04 at 04:22:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:22:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material are used for the insulation between two shells of a helium tank +2025-04-04 at 04:22:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:22:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material was used for the insulation between the two shells of a helium tank +2025-04-04 at 04:22:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:22:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aluminum+helium+insulation+tank +2025-04-04 at 04:22:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:22:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:22:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What insulation material is used for the annular vacuum in helium tanks to prevent contamination and degradation? +2025-04-04 at 04:22:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:22:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: issue with helium tank insulation and cryogenic oxygen tank in Apollo missions +2025-04-04 at 04:22:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:22:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: intended for helium tank insulation material +2025-04-04 at 04:22:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:22:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aluminized+Mylar+helium+tank +2025-04-04 at 04:22:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:22:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:22:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: materials used for annular vacuum in helium or oxygen tanks +2025-04-04 at 04:22:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:22:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 17 cryogenic oxygen and helium tank insulation issue +2025-04-04 at 04:22:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:22:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: material used for helium tank insulation +2025-04-04 at 04:22:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:22:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo helium tank insulation material +2025-04-04 at 04:22:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:22:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:22:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Which material is used for the annular vacuum insulation around helium tanks in the Apollo spacecraft? +2025-04-04 at 04:22:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:22:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo helium tank insulation material +2025-04-04 at 04:22:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:22:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank inner shell insulation material +2025-04-04 at 04:22:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:22:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: stainless steel Mylar insulation helium tank +2025-04-04 at 04:22:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:22:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:22:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo helium tank temperature and pressure issue +2025-04-04 at 04:22:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:22:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aluminized+Mylar+double-walled+helium+tank+annular+region +2025-04-04 at 04:22:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:22:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:22:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo helium tank temperature and pressure screening test +2025-04-04 at 04:22:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:22:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aluminized+Mylar+annular+region+helium+tank +2025-04-04 at 04:22:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:22:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:22:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo helium tank insulation material and pressure rise rates +2025-04-04 at 04:22:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:22:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aluminized+Mylar+annular+region+helium+tank+titanium +2025-04-04 at 04:22:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:22:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:22:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: material+ helium+ tank+ spacecraft+ Apollo +2025-04-04 at 04:22:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:22:43 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:22:43 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:22:43 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, True, True, False] +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_correctness:82 - Student lengths: [296, 32, 150, 319, 278, 1933] +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [16, 16, 16, 16, 16, 16] +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_correctness:84 - Average student length: 501.33 +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 16.00 +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_correctness:86 - Length ratio: 31.33 +2025-04-04 at 04:22:43 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:22:43 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.483 Âą 0.388 +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.67 Âą 2.56 +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 4, 1, 7, 4, 0] +2025-04-04 at 04:22:43 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:22:43 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe most likely cause of the anomaly is a tank-insulation degradation which w...'] +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:22:43 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe most likely cause of the anomaly is a tank-insulation degradation which w...', 'Result 1:\nThe most likely cause of the anomaly is a tank-insulation degradation which w...', 'Result 1:\nThe most likely cause of the anomaly is a tank-insulation degradation which w...', 'Result 1:\nThe most likely cause of the anomaly is a tank-insulation degradation which w...'] +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:22:43 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.533, Perfect scores: 2/6 +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 3.00 +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:22:43 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:22:43 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.438, Max reward: 0.592 +2025-04-04 at 04:22:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:22:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spaccraft initial pitch system +2025-04-04 at 04:22:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The crew reported a pitch-up disturbance torque was exerted on the command module soon after undocking until the beginning of entry. Most of this time, only low-bit-rate telemetry was available and therefore a detailed analysis is impossible. A 20-minute segment of high-bit-rate data was received just prior to entry, and an unaccountable pitch-up torque of 0.00l deg/sec2 was observed. The possible contributing causes for this torque could have been gravity gradients, atmospheric trimming, venting through the umbilical, venting through the tunnel hatch, and a gradual propellant leak. However, none of these is considered to have been a single cause, and either a combination of these causes was present Or some undetermined venting took place. +------ + +2025-04-04 at 04:22:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the initial pitch of a spacecraft? in aerospace engineering +2025-04-04 at 04:22:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ + +2025-04-04 at 04:22:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: initial pitch of Mars Curiosity Rover +2025-04-04 at 04:22:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ + +2025-04-04 at 04:22:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar lunar spacecraft launch pitch +2025-04-04 at 04:22:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:22:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "initial pitch of a spacecraft engine" +2025-04-04 at 04:22:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:22:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft pitch angle +2025-04-04 at 04:22:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:22:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:22:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: initial pitch of a spacecraft in launch state +2025-04-04 at 04:22:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:22:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 spacecraft initial pitch +2025-04-04 at 04:22:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:22:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: initial pitch of the Mars Curiosity Rover at launch +2025-04-04 at 04:22:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:22:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: initial pitch of lunar module +2025-04-04 at 04:22:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:22:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo spacecraft initial pitch angle" +2025-04-04 at 04:22:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:22:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 initial pitch maneuver +2025-04-04 at 04:22:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:22:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:22:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft pitch attitude at liftoff +2025-04-04 at 04:22:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-04 at 04:22:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch inclination or escape velocity of a Mars Curiosity Rover launch +2025-04-04 at 04:22:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ + +2025-04-04 at 04:22:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module initial pitch +2025-04-04 at 04:22:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:22:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo spacecraft pitch attitude at launch" +2025-04-04 at 04:22:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:22:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar orbit insertion maneuvers +2025-04-04 at 04:22:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:22:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:22:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo command and service module S-IVB launch +2025-04-04 at 04:22:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:22:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: initial pitch angle of a Mars rover launch +2025-04-04 at 04:22:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:22:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar module lunar module attitude east launch +2025-04-04 at 04:22:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:22:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo spacecraft launch pitch angle" +2025-04-04 at 04:22:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:22:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 initial pitch maneuver purpose +2025-04-04 at 04:22:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:22:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:22:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: slope of lunar impact trajectory by S-IVB +2025-04-04 at 04:22:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-04 at 04:22:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: launch pitch of a spacecraft +2025-04-04 at 04:22:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:22:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 launch pitch lunar module +2025-04-04 at 04:22:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:22:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo spacecraft launch configuration pitch angle" +2025-04-04 at 04:22:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:22:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Earth orbit to lunar orbit pitch maneuver +2025-04-04 at 04:22:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:22:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:22:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar orbit pitch attitude after S-IVB dump +2025-04-04 at 04:22:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:22:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: initial pitch angle of a Saturn V rocket +2025-04-04 at 04:22:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-04 at 04:22:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module attitude at liftoff +2025-04-04 at 04:22:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-04 at 04:22:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo spacecraft pitch attitude at launch" +2025-04-04 at 04:22:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:22:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Earth's terminator alignment +2025-04-04 at 04:22:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-04 at 04:22:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:23:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electrostatic potentials of a Saturn V launch +2025-04-04 at 04:23:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-04 at 04:23:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module pitch at launch +2025-04-04 at 04:23:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:23:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo S-IVB steering engine initial pitch" +2025-04-04 at 04:23:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:23:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: abort guidance system alignment +2025-04-04 at 04:23:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-04 at 04:23:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:23:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V launch azimuth and initial pitch angle +2025-04-04 at 04:23:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-04 at 04:23:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module pitch angle launch +2025-04-04 at 04:23:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:23:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo S-IC engine pitch angle at T+4.28 seconds" +2025-04-04 at 04:23:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:23:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 post-Earth orbit pitch attitude +2025-04-04 at 04:23:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:23:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:23:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V orientation at launch +2025-04-04 at 04:23:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ + +2025-04-04 at 04:23:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module pitch angle at launch +2025-04-04 at 04:23:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:23:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 S-IC engine pitch angle" +2025-04-04 at 04:23:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:23:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 S-IVB initial separation pitch maneuver +2025-04-04 at 04:23:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:23:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:23:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 S-IVB attitude at T+2 minutes after launch" +2025-04-04 at 04:23:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:23:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB lunar impact attitude +2025-04-04 at 04:23:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-04 at 04:23:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:23:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 S-IVB attitude at 2 minutes post-launch, pitch angle" +2025-04-04 at 04:23:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:23:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB camera attitude +2025-04-04 at 04:23:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:23:09 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:23:09 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:23:09 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_correctness:82 - Student lengths: [251, 277, 1849, 1150, 1406, 1406] +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_correctness:84 - Average student length: 1056.50 +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 4.00 +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_correctness:86 - Length ratio: 264.12 +2025-04-04 at 04:23:09 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_format:228 - Responses ending properly: 6/6 +2025-04-04 at 04:23:09 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.000 Âą 0.000 +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 0.00 Âą 0.00 +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 0, 0, 0, 0] +2025-04-04 at 04:23:09 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The crew reported a pitch-up disturbance torque was exerted on the command module soon after undocking until the beginning of entry. Most of this time, only low-bit-rate telemetry was available and therefore a detailed analysis is impossible. A 20-minute segment of high-bit-rate data was received just prior to entry, and an unaccountable pitch-up torque of 0.00l deg/sec2 was observed. The possible contributing causes for this torque could have been gravity gradients, atmospheric trimming, venting through the umbilical, venting through the tunnel hatch, and a gradual propellant leak. However, none of these is considered to have been a single cause, and either a combination of these causes was present Or some undetermined venting took place. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:23:09 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nCondition Maneuver Second midcourse correction Transearth injection Third mid...', 'Result 1:\nAt lift-off, measured winds, both at the surface and in the region of maximum...', 'Result 1:\nAt lift-off, measured winds, both at the surface and in the region of maximum...', 'Result 1:\nThe discarded S-IVB stage was targeted for a lunar impact of 3 degrees south ...', 'Result 1:\nThe discarded S-IVB stage was targeted for a lunar impact of 3 degrees south ...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...'] +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +The first midcourse correction maneuver, performed at the second option point, was completely nominal. The service propulsion engine was started and stopped on time, and residuals were negligible. In conjunction with this service propulsion maneuver, some differences were noted with respect to the command module simulator. When gimbal motors were turned on, an 8- to l0-ampere increase was noted, with a slightly faster jump than had been seen in the simulator. The major distinction was the fact that fuel cell flowrate indications are barely seen to move, whereas there is a very noticeable change in the simulator. At engine ignition, the ball valve indicators moved slowly to open, but in the simulator, they instantaneously move to open. After turning off the battery bus ties, the battery voltage slowly rose from 32 volts to the open circuit voltage of about 37 volts, whereas in the simulator there is an instantaneous recovery. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ + +2025-04-04 at 04:23:09 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\n1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJEC...', 'Result 1:\n1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJEC...', 'Result 1:\nThe vehicle was launched on an azimuth 90 degrees east of north, and a roll m...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...', 'Result 1:\nAt lift-off, measured winds, both at the surface and in the region of maximum...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nAt lift-off, measured winds, both at the surface and in the region of maximum...', 'Result 1:\nAt lift-off, measured winds, both at the surface and in the region of maximum...'] +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-04 at 04:23:09 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nThe unprecedented powered-down state of the command module required generatio...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nThe unprecedented powered-down state of the command module required generatio...', 'Result 1:\nThe lunar module platform was coarse aligned to the command module platform a...', 'Result 1:\nThe lunar module platform was coarse aligned to the command module platform a...'] +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:23:09 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nThe pilot describing function experiment (T-029) was a success, in that data ...', 'Result 1:\nAt approximately 105 hours, the crew performed a manual descent propulsion ma...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nIgnition and lift-off occurred on schedule. A listing_of major flight plan ev...', 'Result 1:\nIgnition and lift-off occurred on schedule. A listing_of major flight plan ev...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...'] +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-04 at 04:23:09 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAt approximately 105 hours, the crew performed a manual descent propulsion ma...', 'Result 1:\nGuidance system performance was again nominal and there were no significant a...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nThe discarded S-IVB stage was targeted for a lunar impact of 3 degrees south ...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...'] +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 1.0 +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.167 +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:23:09 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.267, Perfect scores: 0/6 +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 3.17 +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 0/6 +2025-04-04 at 04:23:09 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:23:09 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.485, Max reward: 0.750 +2025-04-04 at 04:23:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:23:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electric safety tool left hand interrupter +2025-04-04 at 04:23:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-04 at 04:23:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What electrical safety device escapes from a left-hand circuit interrupter +2025-04-04 at 04:23:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-04 at 04:23:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What electrical component is left behind in a left-hand circuit interrupter +2025-04-04 at 04:23:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-04 at 04:23:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: " Regards to electrical circuit interrupter failure site" +2025-04-04 at 04:23:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-04 at 04:23:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of electrical safety device is commonly referred to as a left-hand circuit interrupter +2025-04-04 at 04:23:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-04 at 04:23:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: left-hand electrical circuit interrupter aircraft +2025-04-04 at 04:23:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-04 at 04:23:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:23:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: causes of gas leakage in command module electrical circuit interrupter +2025-04-04 at 04:23:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ + +2025-04-04 at 04:23:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aerospace electrical safety devices +2025-04-04 at 04:23:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:23:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module vacuum integrity escape gas cause" +2025-04-04 at 04:23:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-04 at 04:23:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: experimental hoáēˇcdn switch or electrical circuit disconnect mechanism used in spacecraft +2025-04-04 at 04:23:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-04 at 04:23:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: function mechanism of command module/service module umbilical guillotine interrupter +2025-04-04 at 04:23:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-04 at 04:23:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:23:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gap in backup ring responsibility +2025-04-04 at 04:23:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +that a constant-g (backup) entry had to be flown. Table l0.3-I lists all the dedicated recovery forces for the Apollo l3 mission. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-04 at 04:23:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module command module electrical circuit interrupter component +2025-04-04 at 04:23:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-04 at 04:23:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module moonwalk gas leak" +2025-04-04 at 04:23:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-04 at 04:23:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft electrical safety mechanism left-hand circuit interrupter +2025-04-04 at 04:23:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-04 at 04:23:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module/service module guillotine propellant +2025-04-04 at 04:23:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-04 at 04:23:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:23:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: forward O-ring valve assembly procedure +2025-04-04 at 04:23:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-04 at 04:23:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module command module electrical connections +2025-04-04 at 04:23:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:23:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lithium hydride gas leak" +2025-04-04 at 04:23:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Evidence indicates that battery 2 may have experienced an electrical fault of some type. The most probable condition is electrolyte leaking from one or more cells and bridging the high-voltage or low-voltage terminal to the battery case (fig. 14-17). This bridging results in water electrolysis and subsequent ignition of the hydrogen and oxygen so generated. The accompanying "explosion" would then blow off or rupture the seal of the battery lid and cause both a thump and venting of the free liquids in the battery case, resulting in "snowflakes." + +Postflight tests have shown the following: +------ +Result 2: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ + +2025-04-04 at 04:23:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gas used in spacecraft safety circuit interrupter +2025-04-04 at 04:23:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-04 at 04:23:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command module/service module propellant oxidizer +2025-04-04 at 04:23:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-04 at 04:23:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:23:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: venting system malfunction Apollo +2025-04-04 at 04:23:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:23:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module descent propulsion firing current surge +2025-04-04 at 04:23:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-04 at 04:23:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lithium hydroxide gas leak" +2025-04-04 at 04:23:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:23:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: equal,  neither helium nor nitrogen, escaped propellant gas from a left-hand electrical circuit interrupter +2025-04-04 at 04:23:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-04 at 04:23:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module reaction control system propellant type +2025-04-04 at 04:23:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-04 at 04:23:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:23:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 systems anomaly +2025-04-04 at 04:23:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-04 at 04:23:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module electrical circuit interrupter left-hand defect +2025-04-04 at 04:23:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-04 at 04:23:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module suit gas leak lithium hydroxide" +2025-04-04 at 04:23:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-04 at 04:23:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: liquid oxygen escaped from left-hand electrical circuit interrupter +2025-04-04 at 04:23:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-04 at 04:23:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system propellant used hydrogen +2025-04-04 at 04:23:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:23:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:23:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 oxygen tank loss +2025-04-04 at 04:23:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-04 at 04:23:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: O-ring failure in hydrogen peroxide system +2025-04-04 at 04:23:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-04 at 04:23:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module oxygen escaped electrical circuit interrupter" +2025-04-04 at 04:23:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:23:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: liquid oxygen pressure relief valve +2025-04-04 at 04:23:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-04 at 04:23:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Oxi di zer chemical +2025-04-04 at 04:23:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-04 at 04:23:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:23:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 oxygen tank failure causes +2025-04-04 at 04:23:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:23:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: O-ring failure left-hand interrupter +2025-04-04 at 04:23:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-04 at 04:23:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: liquid oxygen escaped propellant gas type +2025-04-04 at 04:23:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-04 at 04:23:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: liquid hydrogen reaction control system oxidizer +2025-04-04 at 04:23:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-04 at 04:23:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:23:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: oxygen tank issue Apollo 13 +2025-04-04 at 04:23:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-04 at 04:23:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: left-hand electrical circuit interrupter O-ring failure +2025-04-04 at 04:23:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-04 at 04:23:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cryogenic valve leak +2025-04-04 at 04:23:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-04 at 04:23:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system operation nominal +2025-04-04 at 04:23:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-04 at 04:23:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:23:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: oxygen tank recovery procedures +2025-04-04 at 04:23:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-04 at 04:23:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is left behind when a left-hand circuit interrupter is triggered +2025-04-04 at 04:23:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-04 at 04:23:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: valve type and material used in left-hand electrical circuit interrupter +2025-04-04 at 04:23:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-04 at 04:23:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:23:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what shuts down left-hand side of lunar module/Command Module +2025-04-04 at 04:23:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:23:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: solid-state electrical circuit interrupter +2025-04-04 at 04:23:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-04 at 04:23:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:23:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: liquid oxygen electrical circuit interrupter +2025-04-04 at 04:23:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-04 at 04:23:41 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:23:41 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:23:41 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, False, False, False, False] +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1636, 232, 1902, 173, 1403, 1384] +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [14, 14, 14, 14, 14, 14] +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_correctness:84 - Average student length: 1121.67 +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 14.00 +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_correctness:86 - Length ratio: 80.12 +2025-04-04 at 04:23:41 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:23:41 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.375 Âą 0.425 +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 2.83 Âą 3.72 +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 1, 0, 7, 0, 9] +2025-04-04 at 04:23:41 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +that a constant-g (backup) entry had to be flown. Table l0.3-I lists all the dedicated recovery forces for the Apollo l3 mission. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Evidence indicates that battery 2 may have experienced an electrical fault of some type. The most probable condition is electrolyte leaking from one or more cells and bridging the high-voltage or low-voltage terminal to the battery case (fig. 14-17). This bridging results in water electrolysis and subsequent ignition of the hydrogen and oxygen so generated. The accompanying "explosion" would then blow off or rupture the seal of the battery lid and cause both a thump and venting of the free liquids in the battery case, resulting in "snowflakes." + +Postflight tests have shown the following: +------ +Result 2: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:23:41 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe two interrupters open the electrical circuits about 30 milliseconds befor...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...', 'Result 1:\nEvidence indicates that battery 2 may have experienced an electrical fault of...', 'Result 1:\nIt became apparent that there were insufficient lithium hydroxide cartridges ...', 'Result 1:\nIt became apparent that there were insufficient lithium hydroxide cartridges ...', 'Result 1:\nBased upon an analysis of the interrupter design, its location, and its relat...'] +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 12: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 5.0 +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.833 +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:23:41 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.367, Perfect scores: 1/6 +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 3.17 +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:23:41 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:23:41 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.507, Max reward: 0.808 +2025-04-04 at 04:23:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:23:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what pressure range increase helium tank helium 13 countdown demonstration test +2025-04-04 at 04:23:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:23:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what pressure range helium tank exhibited during rise-rate characteristics during the countdown demonstration test +2025-04-04 at 04:23:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:23:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise-rate characteristics Apollo 13 countdown demonstration test +2025-04-04 at 04:23:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:23:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure range helium tank rise-rate characteristics countdown demonstration test Apollo 13 +2025-04-04 at 04:23:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:23:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:23:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure range helium tank pressure rise rate Apollo 13 countdown demonstration test +2025-04-04 at 04:23:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:23:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: rise-rate characteristics helium tank pressure range 640-900 psia +2025-04-04 at 04:23:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:23:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank pressure range increase rate during Apollo 13 countdown demonstration test +2025-04-04 at 04:23:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:23:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure rise rate helium tank Apollo 13 aircraft leaking helium tank pressure manipulated to 900 psia +2025-04-04 at 04:23:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:23:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:23:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: range of pressure at helium tank abnormal rise rate 14.9 psi/hour +2025-04-04 at 04:23:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:23:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank pressure range at 14.9 psi/hour +2025-04-04 at 04:23:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:23:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Helium tank pressure range start of abnormal rise rate increase during Apollo 13 countdown demonstration test +2025-04-04 at 04:23:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:23:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure rise rate helium tank Apollo 13 manipulated to 900 psia +2025-04-04 at 04:23:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:23:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:23:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: range of pressure helium tank rise rate 14.9 psi/hour +2025-04-04 at 04:23:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:23:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank pressure range above 900 psia 14.9 psi/hour +2025-04-04 at 04:23:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:23:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Helium tank pressure decrease from anomalous rise rate onset +2025-04-04 at 04:23:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-04 at 04:23:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure rise rate helium tank Apollo 13 countdown demonstration test 640 psia +2025-04-04 at 04:23:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:23:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:23:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: compression limit helium tank +2025-04-04 at 04:23:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:23:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Helium tank pressure start of abnormal rise rate anomaly +2025-04-04 at 04:23:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-04 at 04:23:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: initial cold-soak period pressure rise rate helium tank test +2025-04-04 at 04:23:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:23:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:23:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: burst pressure helium tank +2025-04-04 at 04:23:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:23:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Helium tank pressure equivalent to start of abnormal rise rate +2025-04-04 at 04:23:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-04 at 04:23:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: screening test pressure rise rate helium temperature range l3 +2025-04-04 at 04:23:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:23:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:24:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum integrity helium tank +2025-04-04 at 04:24:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:24:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 640 psia helium tank pressure +2025-04-04 at 04:24:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:24:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: initial cold-soak pressure helium tank rise rate +2025-04-04 at 04:24:01 | INFO | src.agent:run_tool_calls:118 - īŋŊīŋŊīŋŊī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:24:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:24:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank contaminants pressure rise rate +2025-04-04 at 04:24:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:24:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: normal and abnormal rise rate helium tank Apollo 13 640 to 900 psia +2025-04-04 at 04:24:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:24:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:24:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: contaminant pressure range helium tank +2025-04-04 at 04:24:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:24:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure range helium tank rise-rate normal and abnormal rise rate 640 psia +2025-04-04 at 04:24:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:24:05 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:24:05 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:24:05 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, True, False, False] +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_correctness:82 - Student lengths: [438, 2013, 728, 306, 523, 1683] +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [24, 24, 24, 24, 24, 24] +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_correctness:84 - Average student length: 948.50 +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 24.00 +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_correctness:86 - Length ratio: 39.52 +2025-04-04 at 04:24:05 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.667, Valid formats: 4.0/6 +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-04 at 04:24:05 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.292 Âą 0.415 +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 1.83 Âą 2.73 +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [4, 0, 0, 7, 0, 0] +2025-04-04 at 04:24:05 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-04 at 04:24:05 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-04 at 04:24:05 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:24:05 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.533, Perfect scores: 2/6 +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 2.67 +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:24:05 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:24:05 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.395, Max reward: 0.744 +2025-04-04 at 04:24:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:24:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: throttle position vs aircraft throttle +2025-04-04 at 04:24:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-04 at 04:24:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the standard throttle position in a typical passenger vehicle and airplane? +2025-04-04 at 04:24:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:24:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: throttle position final + + +2025-04-04 at 04:24:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:24:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Final throttle position in aircraft context +2025-04-04 at 04:24:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: throttle position final Harley-Davidson ​​ electric +2025-04-04 at 04:24:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-04 at 04:24:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aircraft throttle position +2025-04-04 at 04:24:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-04 at 04:24:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:24:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space shuttle throttle position definition +2025-04-04 at 04:24:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-04 at 04:24:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection throttle profile +2025-04-04 at 04:24:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: final transearth injection throttle position +2025-04-04 at 04:24:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection maneuver engine throttle setting +2025-04-04 at 04:24:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +2025-04-04 at 04:24:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:24:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth maneuver throttle position +2025-04-04 at 04:24:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:24:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: shuttle engine throttling profile +2025-04-04 at 04:24:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:24:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (minimum to maximum altitude or minimum throttle position) for a spacecraft maneuver +2025-04-04 at 04:24:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-04 at 04:24:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: final transearth injection steady-state throttle percentage +2025-04-04 at 04:24:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: third to last descent propulsion operation throttle position +2025-04-04 at 04:24:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:24:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: LIOW transit VECH roll and yaw velocity error +2025-04-04 at 04:24:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +With primary guidance available, guidance system steering was manually followed to trim the spacecraft attitudes for the maneuver. Although the displayed attitudes looked favorable in comparison to groundsupplied and out-the-window readings , the primary guidance steering needles read full scale left in roll and yaw (section 6.4). At about 137 hours 40 minutes, the lunar module reaction control system was used to provide a 2.9-ft/sec velocity correction. The maneuver was completed using manual pitch and roll control and abort guidance yaw control in a manner similar to that for the previous midcourse correction. + +8.9.4 Service Module Separation and Photography +------ +Result 2: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ + +2025-04-04 at 04:24:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: throttle position minimum last engine shut down transearth injection +2025-04-04 at 04:24:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:24:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:24:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: saturn v thrust vectoring +2025-04-04 at 04:24:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:24:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: minimum altitude of orbit +2025-04-04 at 04:24:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:24:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: final descent propulsion steady-state throttle percentage +2025-04-04 at 04:24:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:24:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module lunar landing maneuver +2025-04-04 at 04:24:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:24:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection throttle minimum 12 percent vs 37 percent +2025-04-04 at 04:24:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:24:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: saturn v thrust vectoring charging +2025-04-04 at 04:24:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-04 at 04:24:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericynthion altitude for APOLLO 12 translunar injection +2025-04-04 at 04:24:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:24:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection throttle 40% +2025-04-04 at 04:24:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollospacecraft separation mission history +2025-04-04 at 04:24:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:24:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injector final full throttle +2025-04-04 at 04:24:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:24:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:24:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: saturn v electrostatic shielding +2025-04-04 at 04:24:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-04 at 04:24:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: free-return versus non-free-return trajectory +2025-04-04 at 04:24:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ +Result 2: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ + +2025-04-04 at 04:24:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection final throttle position +2025-04-04 at 04:24:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 failure recovery +2025-04-04 at 04:24:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:24:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:24:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 final throttle position +2025-04-04 at 04:24:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:24:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection min throttle percentage +2025-04-04 at 04:24:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 landing recovery +2025-04-04 at 04:24:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:24:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:24:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: throttle position apollo transearth +2025-04-04 at 04:24:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:24:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection throttle 10% +2025-04-04 at 04:24:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module return journey +2025-04-04 at 04:24:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-04 at 04:24:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:24:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth throttle profile +2025-04-04 at 04:24:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection throttle vs 40% +2025-04-04 at 04:24:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar abort +2025-04-04 at 04:24:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-04 at 04:24:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:24:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: min throttle position transearth +2025-04-04 at 04:24:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-04 at 04:24:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 navigation +2025-04-04 at 04:24:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:24:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:24:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 navigation apollo 13 +2025-04-04 at 04:24:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:24:33 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:24:33 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:24:33 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, False, False, False, True] +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_correctness:82 - Student lengths: [2037, 272, 2017, 617, 1956, 393] +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_correctness:84 - Average student length: 1215.33 +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 13.00 +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_correctness:86 - Length ratio: 93.49 +2025-04-04 at 04:24:33 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.833, Valid formats: 5.0/6 +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_format:228 - Responses ending properly: 6/6 +2025-04-04 at 04:24:33 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.317 Âą 0.448 +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 1.83 Âą 2.61 +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 6, 0, 0, 0, 5] +2025-04-04 at 04:24:33 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-04 at 04:24:33 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nAt lift-off, measured winds, both at the surface and in the region of maximum...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...'] +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ +Result 2: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ + +2025-04-04 at 04:24:33 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nvelocity vector onto the local body-centered, horizontal plane, measured posi...', 'Result 1:\nAs on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return t...', 'Result 1:\ndata. Following this maneuver, a series of earth photographs were taken for l...'] +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:33 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nCondition Maneuver Second midcourse correction Transearth injection Third mid...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...'] +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:24:33 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...'] +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +With primary guidance available, guidance system steering was manually followed to trim the spacecraft attitudes for the maneuver. Although the displayed attitudes looked favorable in comparison to groundsupplied and out-the-window readings , the primary guidance steering needles read full scale left in roll and yaw (section 6.4). At about 137 hours 40 minutes, the lunar module reaction control system was used to provide a 2.9-ft/sec velocity correction. The maneuver was completed using manual pitch and roll control and abort guidance yaw control in a manner similar to that for the previous midcourse correction. + +8.9.4 Service Module Separation and Photography +------ +Result 2: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:24:33 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not ...', 'Result 1:\n(section ll.3).\n------\nResult 2:\n+21.0 +3.0\n------\n...', 'Result 1:\nWith primary guidance available, guidance system steering was manually follow...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nThe Mission Control Center and the Manned Space Flight Network provided excel...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nA.2 LUNAR MODULE\n------\nResult 2:\n1.0 SUMMARY\n\nThe Apollo l3 mission, planned...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...'] +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:24:33 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...'] +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 0.0 +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.000 +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.000 +2025-04-04 at 04:24:33 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.733, Perfect scores: 2/6 +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 6/6 +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 4.00 +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:24:33 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:24:33 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.525, Max reward: 0.800 +2025-04-04 at 04:24:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:24:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: trainers pilots aviation training +2025-04-04 at 04:24:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-04 at 04:24:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: flight training section +2025-04-04 at 04:24:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ + +2025-04-04 at 04:24:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aviation pilot training manual or procedures section +2025-04-04 at 04:24:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-04 at 04:24:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training manual +2025-04-04 at 04:24:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-04 at 04:24:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:24:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar mission pilot training apollo +2025-04-04 at 04:24:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:24:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training section +2025-04-04 at 04:24:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ + +2025-04-04 at 04:24:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: types of pilot training +2025-04-04 at 04:24:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-04 at 04:24:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: astronaut training program +2025-04-04 at 04:24:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:24:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:24:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 crew training challenges +2025-04-04 at 04:24:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:24:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training manual section +2025-04-04 at 04:24:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-04 at 04:24:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training aviation +2025-04-04 at 04:24:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-04 at 04:24:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA astronaut training program +2025-04-04 at 04:24:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:24:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:24:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 launch training report +2025-04-04 at 04:24:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:24:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA astronaut training manual +2025-04-04 at 04:24:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:24:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training section +2025-04-04 at 04:24:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ + +2025-04-04 at 04:24:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Astronaut training NASA +2025-04-04 at 04:24:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:24:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:24:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 astronauts training procedures +2025-04-04 at 04:24:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:24:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA astronaut manual 1970 +2025-04-04 at 04:24:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:24:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: section 8.0 pilot training +2025-04-04 at 04:24:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-04 at 04:24:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: US astronaut training +2025-04-04 at 04:24:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:24:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:24:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: _apollo 13 pilot training manual +2025-04-04 at 04:24:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:24:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Apollo 13 pilot training manual +2025-04-04 at 04:24:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:24:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:24:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 pilot crew training schedule +2025-04-04 at 04:24:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:24:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 pilot training geology +2025-04-04 at 04:24:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:24:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:24:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module pilot training +2025-04-04 at 04:24:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:24:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA pilot training geology Apollo 13 +2025-04-04 at 04:24:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:24:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:24:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar pilot training Apollo 13 +2025-04-04 at 04:24:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:24:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:24:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 training procedures commander +2025-04-04 at 04:24:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:24:52 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:24:52 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:24:52 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, True, True] +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1984, 312, 259, 290, 181, 580] +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [3, 3, 3, 3, 3, 3] +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_correctness:84 - Average student length: 601.00 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 3.00 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_correctness:86 - Length ratio: 200.33 +2025-04-04 at 04:24:52 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:24:52 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.475 Âą 0.475 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 3.00 Âą 3.16 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 0, 8, 5, 5] +2025-04-04 at 04:24:52 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:24:52 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nuse of field radios, extravehicular equipment, and assistance from mission co...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +2025-04-04 at 04:24:52 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +2025-04-04 at 04:24:52 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-04 at 04:24:52 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\ndiscussion of the flight and associated crew activities is presented in secti...', 'Result 1:\ndiscussion of the flight and associated crew activities is presented in secti...', 'Result 1:\ndiscussion of the flight and associated crew activities is presented in secti...', 'Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...', 'Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-04 at 04:24:52 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\ndiscussion of the flight and associated crew activities is presented in secti...', 'Result 1:\nuse of field radios, extravehicular equipment, and assistance from mission co...', 'Result 1:\nuse of field radios, extravehicular equipment, and assistance from mission co...', 'Result 1:\ndiscussion of the flight and associated crew activities is presented in secti...', 'Result 1:\ndiscussion of the flight and associated crew activities is presented in secti...'] +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-04 at 04:24:52 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe performance of the abort guidance system and all attitude control aspects...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 0.0 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.000 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.000 +2025-04-04 at 04:24:52 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.433, Perfect scores: 2/6 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 3.33 +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:24:52 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:24:52 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.255, Max reward: 0.475 +2025-04-04 at 04:24:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:24:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: isolation valves testing for isolating functions in spacecraft wiring +2025-04-04 at 04:24:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:24:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: confirm testing procedure for isolation valves on spacecraft control systems +2025-04-04 at 04:24:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:24:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum testing isolation valves on spacecraft electrical systems +2025-04-04 at 04:24:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:24:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft electrical isolation testing validation +2025-04-04 at 04:24:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:24:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spaceworth testing +2025-04-04 at 04:24:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:24:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:25:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control isolation valve improper wiring or connection +2025-04-04 at 04:25:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: confirm isolation valve wiring verification process in space missions +2025-04-04 at 04:25:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:25:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: individual resistance testing for wiring faults +2025-04-04 at 04:25:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-04 at 04:25:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: functional and resistance checks for isolation valves in spacecraft +2025-04-04 at 04:25:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:25:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:25:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: magnetic field induced valve closure malfunction testing +2025-04-04 at 04:25:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-04 at 04:25:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: isolation valve wiring testing methods for proper isolation valve function +2025-04-04 at 04:25:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control isolation valve resistance testing +2025-04-04 at 04:25:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: resistance testing for isolation valve miswiring +2025-04-04 at 04:25:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:25:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: isolation valve wiring isolation for galley space electrical connections +2025-04-04 at 04:25:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:25:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: isolation valve functionality verification beyond functional checks for accurate isolation valve operation +2025-04-04 at 04:25:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reactinol valve resistance testing method +2025-04-04 at 04:25:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: resistance testing for isolation valves at lower voltage levels +2025-04-04 at 04:25:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-04 at 04:25:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:25:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effect of isolation valve wiring designs on specific galley space electrical connections +2025-04-04 at 04:25:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-04 at 04:25:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: inductive testing for isolation valves on critical spacecraft systems +2025-04-04 at 04:25:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: isolation valve resistance testing using functional checks +2025-04-04 at 04:25:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: resistance checking specific to isolation valve wiring +2025-04-04 at 04:25:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:25:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: malfunction of isolation valve valve-lock mechanism +2025-04-04 at 04:25:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-04 at 04:25:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: inductive testing guidelines for isolation valves +2025-04-04 at 04:25:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: resistance testing for isolation valve proper wiring +2025-04-04 at 04:25:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: non-invasive or non-functional testing for isolation valve wiring +2025-04-04 at 04:25:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-04 at 04:25:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:25:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: understand the cause of partial disengagement of locking mechanism +2025-04-04 at 04:25:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-04 at 04:25:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Overrides for isolation valve testing under specific bus voltage conditions +2025-04-04 at 04:25:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: isolation valve testing for proper wiring using magnetic fields +2025-04-04 at 04:25:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: isolation valve testing procedures +2025-04-04 at 04:25:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:25:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mechanical operation of automatic valve locking mechanism +2025-04-04 at 04:25:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-04 at 04:25:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: validation testing for isolation valves beyond functional checks +2025-04-04 at 04:25:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: isolation valve testing with induced magnetic field +2025-04-04 at 04:25:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:25:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: isolation valve wiring verification +2025-04-04 at 04:25:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:25:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: valve locking mechanism design modifications to improve partial travel +2025-04-04 at 04:25:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ + +2025-04-04 at 04:25:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: inductive coupling isolation valve analysis standards +2025-04-04 at 04:25:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: intentional locking or closing mechanism for isolation valves +2025-04-04 at 04:25:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ + +2025-04-04 at 04:25:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: validation of isolation valve circuitry +2025-04-04 at 04:25:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:25:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: resŅ‚иĐŧing valve locking mechanism reÃļrenung +2025-04-04 at 04:25:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-04 at 04:25:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: testing protocols for multi-component spacecraft wiring verification +2025-04-04 at 04:25:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:25:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:25:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: re-tensioning valve locking mechanism based valve operation +2025-04-04 at 04:25:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-04 at 04:25:19 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:25:19 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:25:19 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1777, 1704, 1672, 676, 1909, 238] +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [17, 17, 17, 17, 17, 17] +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_correctness:84 - Average student length: 1329.33 +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 17.00 +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_correctness:86 - Length ratio: 78.20 +2025-04-04 at 04:25:19 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:25:19 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.000 Âą 0.000 +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 0.00 Âą 0.00 +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 0, 0, 0, 0] +2025-04-04 at 04:25:19 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-04 at 04:25:19 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 5.0 +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.833 +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:25:19 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.167, Perfect scores: 0/6 +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 0.50 +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 0/6 +2025-04-04 at 04:25:19 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:25:19 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.482, Max reward: 0.717 +2025-04-04 at 04:25:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:25:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the location of the drain point sensor in a water separator drain tank +2025-04-04 at 04:25:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:25:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the function of a water separator drain tank in a vehicle's cooling system and its typical location? +2025-04-04 at 04:25:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ + +2025-04-04 at 04:25:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: located point water separator drain tank plumbing system +2025-04-04 at 04:25:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-04 at 04:25:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the original location of the drain tank sensing point +2025-04-04 at 04:25:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ + +2025-04-04 at 04:25:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the typical location of a water separator drain tank sensing point in vehicles or equipment? +2025-04-04 at 04:25:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ + +2025-04-04 at 04:25:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: original location of water separator drain tank automotive +2025-04-04 at 04:25:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-04 at 04:25:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:25:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: sensing point water separator drain tank spacecraft +2025-04-04 at 04:25:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ + +2025-04-04 at 04:25:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the original location of the sensing point for the water separator drain tank in the Apollo 7 lunar module's environmental control system? +2025-04-04 at 04:25:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:25:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: surge tank pressure regulator control valve location +2025-04-04 at 04:25:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-04 at 04:25:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: design of one ring fuel cell oxygen tank 2 sensor location +2025-04-04 at 04:25:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-04 at 04:25:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Location of water separator drain tank sensing point in spacecraft or equipment with fuel cell systems +2025-04-04 at 04:25:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-04 at 04:25:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Roving Vehicle ascent stage propulsion system modifications +2025-04-04 at 04:25:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:25:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:25:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: implementation location water separator drain tank spacecraft +2025-04-04 at 04:25:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:25:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the Apollo 7 mission report that contains information about the sensing point for the water separator drain tank in the lunar module's environmental control system? +2025-04-04 at 04:25:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:25:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure regulator control valve location oxygen tank +2025-04-04 at 04:25:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-04 at 04:25:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: one ring water separator drain tank location +2025-04-04 at 04:25:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:25:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Location of the water separator drain tank sensing point in NASA's Space Shuttle +2025-04-04 at 04:25:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ + +2025-04-04 at 04:25:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module water separator drain tank location +2025-04-04 at 04:25:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:25:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:25:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water separator drain tank sensing point location engineering design +2025-04-04 at 04:25:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-04 at 04:25:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the specific Apollo 7 document that contains the description of the water separator drain tank and its sensing point? +2025-04-04 at 04:25:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:25:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: potable water transducer location +2025-04-04 at 04:25:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ + +2025-04-04 at 04:25:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ascension stage shutoff valve forward O-ring installation procedure +2025-04-04 at 04:25:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-04 at 04:25:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Location of water separator drain tank on the Apollo Spacecraft's Life Support System +2025-04-04 at 04:25:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-04 at 04:25:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module lunar surface operations water transfer and pressurization +2025-04-04 at 04:25:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:25:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:25:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water separator drain tank sensing point location +2025-04-04 at 04:25:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:25:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the Apollo 7 mission report that contains information about the sensing point for the water separator drain tank in the lunar module's environmental control system? +2025-04-04 at 04:25:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:25:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: quantity gaging system location suit pressure transducer +2025-04-04 at 04:25:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-04 at 04:25:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ascension stage water separator drain tank +2025-04-04 at 04:25:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-04 at 04:25:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module (CM) water separator drain tank location +2025-04-04 at 04:25:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:25:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module reaction control thrusters water separator drain tank +2025-04-04 at 04:25:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:25:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:25:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water separator drain tank design implementation +2025-04-04 at 04:25:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-04 at 04:25:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Is the Apollo 7 mission report useful in identifying the original sensing point for the water separator drain tank in the lunar module's environmental control system? +2025-04-04 at 04:25:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:25:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: quantity gaging system plumbing location +2025-04-04 at 04:25:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-04 at 04:25:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module water separator drain tank location +2025-04-04 at 04:25:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:25:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module's (CM)bay 4 and service module configuration +2025-04-04 at 04:25:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-04 at 04:25:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module reaction control system procedures +2025-04-04 at 04:25:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:25:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:25:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water separator drain tank sensing point location in spacecraft cooling system +2025-04-04 at 04:25:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:25:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the overall description of the environmental control system's water separator drain tank in the NASA Apollo mission reports? +2025-04-04 at 04:25:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:25:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: quantity gaging system surge tank plumbing location +2025-04-04 at 04:25:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-04 at 04:25:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module water separator drain system location +2025-04-04 at 04:25:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:25:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Service Module bay 4 layout and function +2025-04-04 at 04:25:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-04 at 04:25:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module Service Propulsion System documentation +2025-04-04 at 04:25:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:25:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:25:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How does the water separator drain tank contribute to the overall environmental control system's performance in the lunar module? +2025-04-04 at 04:25:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:25:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: quantity gaging system tank 2 location +2025-04-04 at 04:25:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:25:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module oxygen pressure tunnel drain tank location +2025-04-04 at 04:25:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-04 at 04:25:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Service Module bay 4 configuration and systems affected +2025-04-04 at 04:25:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-04 at 04:25:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module manual or documentation specific component locations +2025-04-04 at 04:25:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-04 at 04:25:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:25:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: oxygen tank 2 quantity gaging system drainage location +2025-04-04 at 04:25:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ + +2025-04-04 at 04:25:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what does "zero water" mean in a metabolic consumption rate +2025-04-04 at 04:25:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-04 at 04:25:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Service Module bay 4 layout and the location of water separator drain tank +2025-04-04 at 04:25:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-04 at 04:25:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Checklist of Lunar Module components +2025-04-04 at 04:25:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-04 at 04:25:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:25:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ascent oxygen tank 2 quantity gaging system bypass valve +2025-04-04 at 04:25:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ + +2025-04-04 at 04:25:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how potable water tank filled +2025-04-04 at 04:25:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ + +2025-04-04 at 04:25:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module separation and reaction control system +2025-04-04 at 04:25:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-04 at 04:25:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module component design or documentation +2025-04-04 at 04:25:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:25:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:25:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module reaction control system and Dagger 1 +2025-04-04 at 04:25:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:25:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module system diagrams or technical drawings +2025-04-04 at 04:25:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-04 at 04:25:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:25:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module subsystem documentation +2025-04-04 at 04:25:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-04 at 04:25:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:25:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module subsystem layout +2025-04-04 at 04:25:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:25:52 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:25:52 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:25:52 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 1/6 answers correct +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, False, False] +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_correctness:82 - Student lengths: [339, 1888, 1659, 1351, 558, 1078] +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [25, 25, 25, 25, 25, 25] +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_correctness:84 - Average student length: 1145.50 +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 25.00 +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_correctness:86 - Length ratio: 45.82 +2025-04-04 at 04:25:52 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.667, Valid formats: 4.0/6 +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_format:228 - Responses ending properly: 6/6 +2025-04-04 at 04:25:52 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.158 Âą 0.354 +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 1.17 Âą 2.61 +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [7, 0, 0, 0, 0, 0] +2025-04-04 at 04:25:52 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ + +2025-04-04 at 04:25:52 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nASHUR Purpose Tests performed Results Environmental Control 109007 To determi...', 'Result 1:\nASHUR Purpose Tests performed Results Environmental Control 109007 To determi...', 'Result 1:\nASHUR Purpose Tests performed Results Environmental Control 109007 To determi...', 'Result 1:\nThe measurement is not essential for flight safety or mission success. The po...', 'Result 1:\nDuring launch the suit pressure transducer reading remained consistent with c...', 'Result 1:\nElectrical shorts in the fan circuit ignited the wire insulation, causing pre...', 'Result 1:\nThe oxygen distribution system operated nominally until deactivation followin...', 'Result 1:\nConsumable usages for the command and service modules prior to the incident w...', 'Result 1:\nThe oxygen distribution system operated nominally until deactivation followin...', 'Result 1:\nDuring the flight, the pressure in the ascent stage oxygen tank 2 increased, ...'] +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ + +2025-04-04 at 04:25:52 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nEfforts to install the tunnel hatch were terminated when the Commander observ...', 'Result 1:\nFollowing recovery of the data, the vehicle had experienced a translation cha...', 'Result 1:\n(fig. 14-20) caused the leakage. When the valve is installed, the forward O-r...', 'Result 1:\n(fig. 14-20) caused the leakage. When the valve is installed, the forward O-r...', 'Result 1:\nThe oxygen distribution system operated nominally until deactivation followin...', 'Result 1:\nHy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 ...', 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...', 'Result 1:\nHy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 ...', 'Result 1:\nThe measurement is not essential for flight safety or mission success. The po...', 'Result 1:\nThe measurement is not essential for flight safety or mission success. The po...'] +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 12: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 13: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:25:52 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.467, Perfect scores: 1/6 +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 5.17 +2025-04-04 at 04:25:52 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:25:52 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:25:53 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.415, Max reward: 0.654 +2025-04-04 at 04:25:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:25:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the standard prelaunch helium tank rise rate? +2025-04-04 at 04:25:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:25:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the prelaunch-standby rise rate of a typical helium tank used in space exploration? +2025-04-04 at 04:25:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:25:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is normal prelaunch-standby rise rate of helium tank +2025-04-04 at 04:25:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:25:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is helium tank buoyancy rate at standard temperature and pressure +2025-04-04 at 04:25:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:25:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the normal prelaunch-standby rise rate of a lunar module's helium tank? +2025-04-04 at 04:25:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:25:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:26:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the helium tank performance of Apollo 11 during launch? +2025-04-04 at 04:26:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 helium tank prelaunch standby rise rate +2025-04-04 at 04:26:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:26:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how high rise rate helium tank varies with temperature +2025-04-04 at 04:26:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:26:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Is the information from the Apollo mission documentation available? Are there specific helium tank reliability data available for the Apollo mission? +2025-04-04 at 04:26:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:26:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:26:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 helium tank prelaunch-standby rise rate +2025-04-04 at 04:26:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:26:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 helium tank hysteresis pressure rise rate +2025-04-04 at 04:26:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rate standard atmospheric conditions +2025-04-04 at 04:26:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:26:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What does the prelaunch-standby rise rate mean in the context of a helium tank's performance, and what NASA documents define this term? +2025-04-04 at 04:26:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:26:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: normal helium tank pressure rise rate +2025-04-04 at 04:26:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank prelaunch rise rate general comparison +2025-04-04 at 04:26:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:26:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rate comparison to 7.9 psi/hour +2025-04-04 at 04:26:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:26:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rates standard atmospheric conditions intact +2025-04-04 at 04:26:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:26:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:26:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium rise rate empty tank standard conditions +2025-04-04 at 04:26:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:26:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rate standard atmospheric pressure and temperature +2025-04-04 at 04:26:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:26:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:26:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rate standard atmospheric pressure and temperature +2025-04-04 at 04:26:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:26:12 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:26:12 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:26:12 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, True, False, False, True] +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.67 +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_correctness:82 - Student lengths: [264, 301, 250, 1683, 838, 87] +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_correctness:84 - Average student length: 570.50 +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 12.00 +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_correctness:86 - Length ratio: 47.54 +2025-04-04 at 04:26:12 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.667, Valid formats: 4.0/6 +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:26:12 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.408 Âą 0.318 +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 1.83 Âą 1.57 +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [1, 3, 4, 0, 0, 3] +2025-04-04 at 04:26:12 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-04 at 04:26:12 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 5.0 +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.833 +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.373 +2025-04-04 at 04:26:12 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.467, Perfect scores: 2/6 +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 2.17 +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:26:12 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:26:12 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.438, Max reward: 0.754 +2025-04-04 at 04:26:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:26:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: irregularities in financial markets +2025-04-04 at 04:26:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ +Result 2: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ + +2025-04-04 at 04:26:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: anomalies in data statistics vs statistical anomalies +2025-04-04 at 04:26:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-04 at 04:26:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the major medical anomalies discussed in clinical literature? +2025-04-04 at 04:26:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ + +2025-04-04 at 04:26:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: anomalies in an unnamed entity's data analysis +2025-04-04 at 04:26:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:26:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: anomalies data location +2025-04-04 at 04:26:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:26:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: anomalies in science +2025-04-04 at 04:26:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-04 at 04:26:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:26:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 1 saturn ib anomaly command module +2025-04-04 at 04:26:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:26:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 reaction control systems data +2025-04-04 at 04:26:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:26:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Spacecraft Electrical Anomalies +2025-04-04 at 04:26:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:26:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth lunar module systems performance data +2025-04-04 at 04:26:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo mission launch site anomalies data +2025-04-04 at 04:26:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-04 at 04:26:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 17 earth orbit insertion anomalies +2025-04-04 at 04:26:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:26:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:26:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system deactivation apollo 1 +2025-04-04 at 04:26:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:26:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system performance apollo missions +2025-04-04 at 04:26:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:26:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 Lightning Strike Analysis +2025-04-04 at 04:26:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-04 at 04:26:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module Ascent and Descent Propulsion Systems performance evaluations +2025-04-04 at 04:26:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 launch site 4/5 +2025-04-04 at 04:26:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:26:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 17 navigation anomalies during transearth injection +2025-04-04 at 04:26:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:26:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:26:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system postflight evaluation +2025-04-04 at 04:26:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 postflight reaction control system tests +2025-04-04 at 04:26:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:26:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Spacecraft Electrostatic Discharge Study +2025-04-04 at 04:26:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-04 at 04:26:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module Ascent and Descent Propulsion System errors or issues +2025-04-04 at 04:26:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:26:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: geology field trip apollo 13 training houston +2025-04-04 at 04:26:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:26:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 17 transearth injection attitude control anomalies +2025-04-04 at 04:26:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:26:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:26:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 1 reaction control system final flight evaluation +2025-04-04 at 04:26:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 command module postflight reaction control system tests table c-1 +2025-04-04 at 04:26:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Apollo Lightning Strike Research +2025-04-04 at 04:26:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-04 at 04:26:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 spacecraft launch cape kennedy +2025-04-04 at 04:26:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:26:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 17 transearth injection trajectory control issues +2025-04-04 at 04:26:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-04 at 04:26:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:26:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 1 command module reaction control system evaluation +2025-04-04 at 04:26:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:26:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 table c-1 command module +2025-04-04 at 04:26:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:26:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Lightning-Induced Electrical Charging on Rockets +2025-04-04 at 04:26:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:26:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 launch site location cape kennedy +2025-04-04 at 04:26:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:26:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 17 transearth injection guidance errors corrections +2025-04-04 at 04:26:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-04 at 04:26:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:26:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system deactivation apollo 1 +2025-04-04 at 04:26:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:26:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 reaction control system table c-i +2025-04-04 at 04:26:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:26:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Why was there a discrepancy between field-change and sferics detector readings and actual lightning strike? +2025-04-04 at 04:26:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-04 at 04:26:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: site 5 location cape kennedy +2025-04-04 at 04:26:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-04 at 04:26:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 17 navigation system guidance errors and corrections during transearth injection +2025-04-04 at 04:26:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:26:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:26:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system evaluation +2025-04-04 at 04:26:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apolo 11 table c-1 command module +2025-04-04 at 04:26:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:26:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Electric Field Meter Placement +2025-04-04 at 04:26:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:26:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: site reference cape kennedy apollo 13 +2025-04-04 at 04:26:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-04 at 04:26:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 17 manual attitude adjustments for transearth injection +2025-04-04 at 04:26:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:26:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:26:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 1 vs lunar module reaction control system evaluation +2025-04-04 at 04:26:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 command module data availability table d-i +2025-04-04 at 04:26:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:26:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Mission Training Challenges +2025-04-04 at 04:26:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:26:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:26:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system evaluations +2025-04-04 at 04:26:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 table d-i +2025-04-04 at 04:26:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:26:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Critical Systems Management +2025-04-04 at 04:26:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:26:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:26:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo lunar module reaction control system evaluation cancellation +2025-04-04 at 04:26:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 reactions control system performance +2025-04-04 at 04:26:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:47 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:26:47 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:26:47 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, True, False] +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_correctness:82 - Student lengths: [863, 863, 1985, 375, 357, 2009] +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [22, 22, 22, 22, 22, 22] +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_correctness:84 - Average student length: 1075.33 +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 22.00 +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_correctness:86 - Length ratio: 48.88 +2025-04-04 at 04:26:47 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-04 at 04:26:47 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.158 Âą 0.354 +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 1.33 Âą 2.98 +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 0, 0, 8, 0] +2025-04-04 at 04:26:47 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ +Result 2: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:47 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\ndata. Following this maneuver, a series of earth photographs were taken for l...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...'] +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:47 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nA.1 COMMAND AND SERVICE MODULES\n------\nResult 2:\nTables D-I and D-II are summ...', 'Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...'] +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-04 at 04:26:47 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nAlthough the standard format was followed during the deactivation and postrec...', 'Result 1:\nAs a result of the electrical disturbances experienced during the Apollo l2 l...', 'Result 1:\nThe field-change and sferics detectors at site 5 gave no indication of any li...', 'Result 1:\nField meter records indicate the Apollo l3 vehicle carried aloft a net positi...', 'Result 1:\nAs a result of the electrical disturbances experienced during the Apollo l2 l...', 'Result 1:\nField meter records indicate the Apollo l3 vehicle carried aloft a net positi...', 'Result 1:\nThe field-change and sferics detectors at site 5 gave no indication of any li...', 'Result 1:\nThe field-change and sferics detectors at site 5 gave no indication of any li...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-04 at 04:26:47 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...'] +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-04 at 04:26:47 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nAt site 6, the record was similar to that for site 7 with an initial positive...', 'Result 1:\nThe field-change and sferics detectors at site 5 gave no indication of any li...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nWhite Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat r...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-04 at 04:26:47 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nEvent Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product o...', 'Result 1:\nAt approximately 105 hours, the crew performed a manual descent propulsion ma...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nInitial outside observations through the lunar module windows indicated that ...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...'] +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 0.0 +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.000 +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.000 +2025-04-04 at 04:26:47 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.533, Perfect scores: 2/6 +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 5.33 +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:26:47 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:26:47 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.525, Max reward: 0.767 +2025-04-04 at 04:26:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:26:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias update value +2025-04-04 at 04:26:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:26:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: accelerometer Z-axis bias update +2025-04-04 at 04:26:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:26:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: update Z-axis accelerometer bias +2025-04-04 at 04:26:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-04 at 04:26:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias update +2025-04-04 at 04:26:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:26:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Z-axis accelerometer bias update after measurement" +2025-04-04 at 04:26:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:26:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:26:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias after update +2025-04-04 at 04:26:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:26:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: robotic ascent abort guidance system Z-axis accelerometer bias +2025-04-04 at 04:26:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Abort guidance system performance was nominal. No instrument calibrations or compensation updates were performed. Uncompensated accelerometer biases and gyro drifts remained within normal operating limits even though heater power was removed from the abort sensor assembly for most of the flight to conserve electrical power. At times, the sensor package temperature was as low as 37? F. + +Accelerometer bias shifts associated with the 30-day and 3-day requirements were well within specification. Table 6.4-II contains preflight calibration histories for the initial components of the abort gui dance system. + +TABLE 6.4-II.- ABORT GUIDANCE SYSTEM PREINSTALLATION CALIBRATION DATA +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:26:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: new z-axis accelerometer bias in reaction control after update +2025-04-04 at 04:26:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-04 at 04:26:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Z-axis accelerometer bias update value +2025-04-04 at 04:26:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:26:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:26:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias measurement after successful flight +2025-04-04 at 04:26:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-04 at 04:26:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: robotic ascent abort guidance system accelerometer bias update +2025-04-04 at 04:26:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Abort guidance system performance was nominal. No instrument calibrations or compensation updates were performed. Uncompensated accelerometer biases and gyro drifts remained within normal operating limits even though heater power was removed from the abort sensor assembly for most of the flight to conserve electrical power. At times, the sensor package temperature was as low as 37? F. + +Accelerometer bias shifts associated with the 30-day and 3-day requirements were well within specification. Table 6.4-II contains preflight calibration histories for the initial components of the abort gui dance system. + +TABLE 6.4-II.- ABORT GUIDANCE SYSTEM PREINSTALLATION CALIBRATION DATA +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:26:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system new z-axis accelerometer bias update +2025-04-04 at 04:26:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-04 at 04:26:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Z-axis accelerometer bias delta value +2025-04-04 at 04:26:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:26:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:27:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias value after successful flight +2025-04-04 at 04:27:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-04 at 04:27:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: robotic ascent abort guidance system Z-axis accelerometer bias shift +2025-04-04 at 04:27:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Abort guidance system performance was nominal. No instrument calibrations or compensation updates were performed. Uncompensated accelerometer biases and gyro drifts remained within normal operating limits even though heater power was removed from the abort sensor assembly for most of the flight to conserve electrical power. At times, the sensor package temperature was as low as 37? F. + +Accelerometer bias shifts associated with the 30-day and 3-day requirements were well within specification. Table 6.4-II contains preflight calibration histories for the initial components of the abort gui dance system. + +TABLE 6.4-II.- ABORT GUIDANCE SYSTEM PREINSTALLATION CALIBRATION DATA +------ +Result 2: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ + +2025-04-04 at 04:27:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control z-axis accelerometer bias update +2025-04-04 at 04:27:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-04 at 04:27:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:27:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias after successful flight +2025-04-04 at 04:27:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-04 at 04:27:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Z axis accelerometer bias shift during robotic ascent +2025-04-04 at 04:27:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-04 at 04:27:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system updated accelerometer bias values +2025-04-04 at 04:27:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-04 at 04:27:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:27:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: new z-axis accelerometer bias value +2025-04-04 at 04:27:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:27:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: robotic ascent abort guidance system Y-axis Z-axis accelerometer bias +2025-04-04 at 04:27:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Abort guidance system performance was nominal. No instrument calibrations or compensation updates were performed. Uncompensated accelerometer biases and gyro drifts remained within normal operating limits even though heater power was removed from the abort sensor assembly for most of the flight to conserve electrical power. At times, the sensor package temperature was as low as 37? F. + +Accelerometer bias shifts associated with the 30-day and 3-day requirements were well within specification. Table 6.4-II contains preflight calibration histories for the initial components of the abort gui dance system. + +TABLE 6.4-II.- ABORT GUIDANCE SYSTEM PREINSTALLATION CALIBRATION DATA +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:27:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias changes during reaction control +2025-04-04 at 04:27:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-04 at 04:27:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:27:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (Z-axis accelerometer bias) robotic ascent abort guidance system +2025-04-04 at 04:27:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Abort guidance system performance was nominal. No instrument calibrations or compensation updates were performed. Uncompensated accelerometer biases and gyro drifts remained within normal operating limits even though heater power was removed from the abort sensor assembly for most of the flight to conserve electrical power. At times, the sensor package temperature was as low as 37? F. + +Accelerometer bias shifts associated with the 30-day and 3-day requirements were well within specification. Table 6.4-II contains preflight calibration histories for the initial components of the abort gui dance system. + +TABLE 6.4-II.- ABORT GUIDANCE SYSTEM PREINSTALLATION CALIBRATION DATA +------ +Result 2: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ + +2025-04-04 at 04:27:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:27:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (Z-axis accelerometer bias change during robotic ascent) +2025-04-04 at 04:27:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-04 at 04:27:08 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:27:08 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:27:08 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_correctness:82 - Student lengths: [276, 1806, 1806, 974, 1854, 284] +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [44, 44, 44, 44, 44, 44] +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_correctness:84 - Average student length: 1166.67 +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 44.00 +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_correctness:86 - Length ratio: 26.52 +2025-04-04 at 04:27:08 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-04 at 04:27:08 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.000 Âą 0.000 +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 0.00 Âą 0.00 +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 0, 0, 0, 0, 0] +2025-04-04 at 04:27:08 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-04 at 04:27:08 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:27:08 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...'] +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Abort guidance system performance was nominal. No instrument calibrations or compensation updates were performed. Uncompensated accelerometer biases and gyro drifts remained within normal operating limits even though heater power was removed from the abort sensor assembly for most of the flight to conserve electrical power. At times, the sensor package temperature was as low as 37? F. + +Accelerometer bias shifts associated with the 30-day and 3-day requirements were well within specification. Table 6.4-II contains preflight calibration histories for the initial components of the abort gui dance system. + +TABLE 6.4-II.- ABORT GUIDANCE SYSTEM PREINSTALLATION CALIBRATION DATA +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Abort guidance system performance was nominal. No instrument calibrations or compensation updates were performed. Uncompensated accelerometer biases and gyro drifts remained within normal operating limits even though heater power was removed from the abort sensor assembly for most of the flight to conserve electrical power. At times, the sensor package temperature was as low as 37? F. + +Accelerometer bias shifts associated with the 30-day and 3-day requirements were well within specification. Table 6.4-II contains preflight calibration histories for the initial components of the abort gui dance system. + +TABLE 6.4-II.- ABORT GUIDANCE SYSTEM PREINSTALLATION CALIBRATION DATA +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Abort guidance system performance was nominal. No instrument calibrations or compensation updates were performed. Uncompensated accelerometer biases and gyro drifts remained within normal operating limits even though heater power was removed from the abort sensor assembly for most of the flight to conserve electrical power. At times, the sensor package temperature was as low as 37? F. + +Accelerometer bias shifts associated with the 30-day and 3-day requirements were well within specification. Table 6.4-II contains preflight calibration histories for the initial components of the abort gui dance system. + +TABLE 6.4-II.- ABORT GUIDANCE SYSTEM PREINSTALLATION CALIBRATION DATA +------ +Result 2: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Abort guidance system performance was nominal. No instrument calibrations or compensation updates were performed. Uncompensated accelerometer biases and gyro drifts remained within normal operating limits even though heater power was removed from the abort sensor assembly for most of the flight to conserve electrical power. At times, the sensor package temperature was as low as 37? F. + +Accelerometer bias shifts associated with the 30-day and 3-day requirements were well within specification. Table 6.4-II contains preflight calibration histories for the initial components of the abort gui dance system. + +TABLE 6.4-II.- ABORT GUIDANCE SYSTEM PREINSTALLATION CALIBRATION DATA +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Abort guidance system performance was nominal. No instrument calibrations or compensation updates were performed. Uncompensated accelerometer biases and gyro drifts remained within normal operating limits even though heater power was removed from the abort sensor assembly for most of the flight to conserve electrical power. At times, the sensor package temperature was as low as 37? F. + +Accelerometer bias shifts associated with the 30-day and 3-day requirements were well within specification. Table 6.4-II contains preflight calibration histories for the initial components of the abort gui dance system. + +TABLE 6.4-II.- ABORT GUIDANCE SYSTEM PREINSTALLATION CALIBRATION DATA +------ +Result 2: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-04 at 04:27:08 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAbort guidance system performance was nominal. No instrument calibrations or ...', 'Result 1:\nAbort guidance system performance was nominal. No instrument calibrations or ...', 'Result 1:\nAbort guidance system performance was nominal. No instrument calibrations or ...', 'Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\nAbort guidance system performance was nominal. No instrument calibrations or ...', 'Result 1:\nAbort guidance system performance was nominal. No instrument calibrations or ...', 'Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...'] +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-04 at 04:27:08 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...'] +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-04 at 04:27:08 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...'] +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-04 at 04:27:08 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...'] +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 0.0 +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.000 +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.000 +2025-04-04 at 04:27:08 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.367, Perfect scores: 1/6 +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 3/6 +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 2.17 +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:27:08 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:27:08 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.258, Max reward: 0.781 +2025-04-04 at 04:27:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:27:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel cell power system average current during ISS mission +2025-04-04 at 04:27:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-04 at 04:27:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: artemis mission sls space launch system fuel cell power system average current +2025-04-04 at 04:27:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:27:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the average current of fuel cells during the space shuttle missionSTS-71B +2025-04-04 at 04:27:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-04 at 04:27:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel cell efficiency NASA International Space Station mission +2025-04-04 at 04:27:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-04 at 04:27:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:27:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel cell current efficiency +2025-04-04 at 04:27:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-04 at 04:27:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: artemis mission fuel cell electrical power system average current bus voltage +2025-04-04 at 04:27:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-04 at 04:27:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the 24 ampere value of current for the space shuttle fuel cells +2025-04-04 at 04:27:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:27:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space shuttle main fuel cell current +2025-04-04 at 04:27:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-04 at 04:27:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:27:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: artemis mission fuel cell current overload caution +2025-04-04 at 04:27:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-04 at 04:27:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is 24 ampere current in ampere-hours (ah) +2025-04-04 at 04:27:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-04 at 04:27:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space shuttle main bus a fuel cell current and main bus b fuel cell current +2025-04-04 at 04:27:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:27:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:27:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: artemis mission fuel cell 3 current +2025-04-04 at 04:27:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:27:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 24 amperes per ah +2025-04-04 at 04:27:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:27:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:27:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: artemis mission fuel cell 2 current +2025-04-04 at 04:27:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:27:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the unit of kWh in ampere-hours (ah) +2025-04-04 at 04:27:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:27:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:27:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: artemis mission fuel cell current fuel cell 2 operation +2025-04-04 at 04:27:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:27:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how many ampere-hours is 120 kWh in +2025-04-04 at 04:27:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:27:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:27:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: artemis mission fuel cell 2 current +2025-04-04 at 04:27:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:27:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: calculate the ratio of 24 amperes per ah +2025-04-04 at 04:27:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +-1.2 40.4 +0.4 +------ + +2025-04-04 at 04:27:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:27:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: artemis mission fuel cell 2 operational parameters +2025-04-04 at 04:27:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal. + +During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l. +------ + +2025-04-04 at 04:27:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: generate a relationship between 24 amperes and ampere-hours +2025-04-04 at 04:27:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:27:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:27:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: artemis mission fuel cell 2 current +2025-04-04 at 04:27:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:27:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how long in hours is 410 ah at 24 amperes +2025-04-04 at 04:27:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:27:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:27:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: artemis mission fuel cell 2 operation details +2025-04-04 at 04:27:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:27:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is 24 amperes sustained for 410 ah +2025-04-04 at 04:27:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:27:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:27:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: artemis mission fuel cell 2 startup current +2025-04-04 at 04:27:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-04 at 04:27:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: calculate how long is 410 ah at 24 amperes +2025-04-04 at 04:27:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +-1.2 40.4 +0.4 +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:27:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:27:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: artemis mission fuel cell 2 crisis current +2025-04-04 at 04:27:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:27:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: time to sustain 24 amperes for 410 ampere hours +2025-04-04 at 04:27:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:27:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:27:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: time to sustain 24 amperes for 410 ah +2025-04-04 at 04:27:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:27:37 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:27:37 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:27:37 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, False, True] +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_correctness:82 - Student lengths: [167, 352, 1142, 1198, 273, 171] +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_correctness:84 - Average student length: 550.50 +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_correctness:86 - Length ratio: 55.05 +2025-04-04 at 04:27:37 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-04 at 04:27:37 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.192 Âą 0.274 +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 0.83 Âą 1.21 +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [2, 0, 0, 0, 0, 3] +2025-04-04 at 04:27:37 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-04 at 04:27:37 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal. + +During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l. +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 12: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +-1.2 40.4 +0.4 +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +-1.2 40.4 +0.4 +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 12: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 13: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-04 at 04:27:37 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 4.0 +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.667 +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:27:37 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.400, Perfect scores: 1/6 +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 3.67 +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 1/6 +2025-04-04 at 04:27:37 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:27:37 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.326, Max reward: 0.716 +2025-04-04 at 04:27:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:27:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature cold soak system electronics aerospace 7 hours +2025-04-04 at 04:27:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-04 at 04:27:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Temperature in cryogenic storage chamber +2025-04-04 at 04:27:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-04 at 04:27:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the temperature typically used to cold soak semiconductor device testing +2025-04-04 at 04:27:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:27:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What standards specify the temperature requirements for cold soaking of electronic components? +2025-04-04 at 04:27:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:27:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature cold soak 7 hours electronics +2025-04-04 at 04:27:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-04 at 04:27:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature cold soak 7 hours +2025-04-04 at 04:27:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:27:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:27:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Temperature of nitrogen or Helium in cold storage from space missions +2025-04-04 at 04:27:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:27:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the typical conditions of a cold soak for semiconductor testing +2025-04-04 at 04:27:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:27:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature control and cabin management control systems for Apollo spacecraft +2025-04-04 at 04:27:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:27:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module command module temperature cold soak 7 hours +2025-04-04 at 04:27:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-04 at 04:27:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:27:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Cryogenic control of helium temperature +2025-04-04 at 04:27:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:27:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the temperature of the system after 7 hours of cold soak +2025-04-04 at 04:27:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-04 at 04:27:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: postflight testing Apollo spacecraft temperature control systems +2025-04-04 at 04:27:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:27:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module temperature cold soak 7 hours +2025-04-04 at 04:27:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-04 at 04:27:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:27:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Temperature threshold for helium condensation +2025-04-04 at 04:27:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:27:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command module cold soak temperature after 7 hours +2025-04-04 at 04:27:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-04 at 04:27:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Charged Particle Lunar Environment Experiment Apollo 11 +2025-04-04 at 04:27:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:27:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:27:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Vaporization temperature of helium in cryogenic storage +2025-04-04 at 04:27:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:27:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (lases with command module temperature) +2025-04-04 at 04:27:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-04 at 04:27:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Charged Particle Lunar Environment Experiment cold cathode gage experiment temperature control +2025-04-04 at 04:27:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:27:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:27:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Critical temperature for helium condensation +2025-04-04 at 04:27:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:27:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (command module helium temperature after cold soak for 7 hours) +2025-04-04 at 04:27:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:27:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: details of post Apollo 17 lunar surface drill hole sensors +2025-04-04 at 04:27:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:27:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:27:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Helium contaminants and tank pressure +2025-04-04 at 04:27:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:27:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module helium temperature during cold soak +2025-04-04 at 04:27:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:27:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 17 cold soak temperature +2025-04-04 at 04:27:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:27:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-04 at 04:28:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Vaporization temperature of hydrogen at -70°C +2025-04-04 at 04:28:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:28:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module descent engine cold soak +2025-04-04 at 04:28:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-04 at 04:28:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:28:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Vapourization temperature range for hydrogen at -70°C +2025-04-04 at 04:28:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:28:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature management Apollo 13 cold soaked lunar landing temperature +2025-04-04 at 04:28:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-04 at 04:28:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:28:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Vaporization temperature range for 10^-6 pounds of hydrogen +2025-04-04 at 04:28:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-04 at 04:28:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: conditions during cold soak Apollo 13 lunar module +2025-04-04 at 04:28:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-04 at 04:28:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:28:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: battery drain power Apollo 13 +2025-04-04 at 04:28:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:28:05 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:28:05 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:28:05 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, False, False, False] +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_correctness:82 - Student lengths: [292, 1903, 172, 1983, 713, 205] +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_correctness:84 - Average student length: 878.00 +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 5.00 +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_correctness:86 - Length ratio: 175.60 +2025-04-04 at 04:28:05 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-04 at 04:28:05 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.237 Âą 0.363 +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 5.33 Âą 9.16 +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 1/6 +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 25, 7, 0, 0, 0] +2025-04-04 at 04:28:05 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-04 at 04:28:05 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nDuring the period when the command module was powered down, the cabin tempera...'] +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ +Result 2: +Consumable quantities in the cryogenic storage system are discussed in section 7.l. + +5.4 COMMUNICATIONS EQUIPMENT + +The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good. + +Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4. +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-04 at 04:28:05 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nCryogenic storage system operation was satisfactory until 46:40:09, when the ...', 'Result 1:\nThe command module reaction control system helium pressures and temperatures ...', 'Result 1:\nThe command module reaction control system helium pressures and temperatures ...', 'Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...', 'Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...', 'Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...', 'Result 1:\nFigure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure.\n...', 'Result 1:\nThe command module reaction control system helium pressures and temperatures ...', 'Result 1:\nThe command module reaction control system helium pressures and temperatures ...', 'Result 1:\nThe most likely cause of the anomaly is a tank-insulation degradation which w...'] +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:28:05 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...', 'Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...', 'Result 1:\nDuring the period when the command module was powered down, the cabin tempera...', 'Result 1:\nDuring the period when the command module was powered down, the cabin tempera...', 'Result 1:\nThe command module reaction control system helium pressures and temperatures ...', 'Result 1:\nThe command module reaction control system helium pressures and temperatures ...', 'Result 1:\nThe command module reaction control system helium pressures and temperatures ...'] +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 11: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-04 at 04:28:05 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nDuring the period when the command module was powered down, the cabin tempera...', 'Result 1:\nTo establish a passive thermal control mode during initial transearth coast, ...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nThe charged particle lunar environment experiment was designed to measure the...', 'Result 1:\nThe charged particle lunar environment experiment was designed to measure the...', 'Result 1:\nThe experiment consisted of an electronics package and two sensor probes whic...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...'] +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-04 at 04:28:05 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nDuring the period when the command module was powered down, the cabin tempera...', 'Result 1:\nEquipment transfer and stowage in both the command module and lunar module wa...', 'Result 1:\nEquipment transfer and stowage in both the command module and lunar module wa...'] +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-04 at 04:28:05 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nDuring the period when the command module was powered down, the cabin tempera...'] +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 0.0 +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.000 +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.000 +2025-04-04 at 04:28:05 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.533, Perfect scores: 2/6 +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 5/6 +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 4.33 +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:28:05 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:28:05 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.423, Max reward: 0.667 +2025-04-04 at 04:28:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:28:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the first acceleration peak during the guided entry of the Apollo missions +2025-04-04 at 04:28:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:28:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the first acceleration peak during the guided entry for apollo 11 +2025-04-04 at 04:28:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:28:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the first acceleration peak during the guided entry of Apollo 11? +2025-04-04 at 04:28:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:28:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the first acceleration peak during spacecraft re-entry? +2025-04-04 at 04:28:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ + +2025-04-04 at 04:28:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Portal acceleration peak" or Portal game info +2025-04-04 at 04:28:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ + +2025-04-04 at 04:28:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-04 at 04:28:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first acceleration peak during guided entry of Apollo 12 mission +2025-04-04 at 04:28:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:28:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: translunar injection acceleration peak during guided entry +2025-04-04 at 04:28:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:28:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: acceleration peak during Apollo 11 guided entry +2025-04-04 at 04:28:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:28:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: acceleration peak first guided entry +2025-04-04 at 04:28:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:28:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-04 at 04:28:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first acceleration peak during Apollo 12 guided entry +2025-04-04 at 04:28:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:28:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first translunar injection acceleration +2025-04-04 at 04:28:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-04 at 04:28:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:28:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first acceleration during guided entry of Apollo 12 mission +2025-04-04 at 04:28:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-04 at 04:28:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: translunar first midcourse correction S-IVB ignition acceleration +2025-04-04 at 04:28:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-04 at 04:28:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:28:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: acceleration pattern during impact phase of space mission +2025-04-04 at 04:28:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:28:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB second midcourse correction translunar acceleration +2025-04-04 at 04:28:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:28:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:28:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar landing acceleration peak descent phase +2025-04-04 at 04:28:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:28:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB translunar injection velocity change +2025-04-04 at 04:28:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:28:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:28:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar landing entry and descent phase guidance +2025-04-04 at 04:28:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:28:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB translunar injection velocity acceleration change +2025-04-04 at 04:28:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:28:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-04 at 04:28:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: manually controlled entry alignment during Apollo 13 descent +2025-04-04 at 04:28:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:28:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:28:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar landing acceleration peak descent phase +2025-04-04 at 04:28:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:28:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-04 at 04:28:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first transearth midcourse correction lunar landing acceleration +2025-04-04 at 04:28:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:28:25 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-04 at 04:28:25 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-04 at 04:28:25 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, True] +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1714, 198, 172, 435, 434, 315] +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [16, 16, 16, 16, 16, 16] +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_correctness:84 - Average student length: 544.67 +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 16.00 +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_correctness:86 - Length ratio: 34.04 +2025-04-04 at 04:28:25 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-04 at 04:28:25 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_retry:330 - Retry behavior rewards: 0.242 Âą 0.366 +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_retry:331 - Search tags per completion: 1.50 Âą 2.57 +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_retry:332 - Violations (>1 search per message): 0/6 +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_retry:333 - Search counts distribution: [0, 7, 0, 0, 0, 2] +2025-04-04 at 04:28:25 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 8: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 9: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 10: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-04 at 04:28:25 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nAt lift-off, measured winds, both at the surface and in the region of maximum...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nAt approximately 105 hours, the crew performed a manual descent propulsion ma...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...'] +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 3: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 4: Result 1: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 5: Result 1: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 6: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 7: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-04 at 04:28:25 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nTranslunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitud...', 'Result 1:\nTranslunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitud...', 'Result 1:\ncrewmen noted the small change in acceleration caused by the mixture ratio sh...', 'Result 1:\ncrewmen noted the small change in acceleration caused by the mixture ratio sh...'] +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-04 at 04:28:25 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...'] +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +2025-04-04 at 04:28:25 | WARNING | src.rewards:reward_em_chunk:392 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:384 - 📝 Ground Truth Chunk: The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 1: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:386 - 🔍 Searched Chunk 2: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:402 - Chunk Query Rewards Summary: +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:403 - Total prompts: 6 +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:404 - Correct matches: 2.0 +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:405 - Average reward: 0.333 +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_em_chunk:406 - Reward std: 0.471 +2025-04-04 at 04:28:25 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_search_strategy:630 - Search strategy metrics - Mean: 0.433, Perfect scores: 1/6 +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_search_strategy:633 - Initial searches: 4/6 +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_search_strategy:634 - Average info processing steps: 2.33 +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_search_strategy:635 - Final synthesis rate: 2/6 +2025-04-04 at 04:28:25 | INFO | src.rewards:log_chat_state:855 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-04 at 04:28:25 | INFO | src.rewards:reward_search_diversity:811 - Search diversity metrics - Mean reward: 0.383, Max reward: 0.759 +2025-04-04 at 04:28:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-04 at 04:28:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were S-II staging and S-IVB ignition times for Saturn V mission +2025-04-04 at 04:28:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:28:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo s-ii s-ivb ignition timeline +2025-04-04 at 04:28:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-04 at 04:28:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 s ii ignition time +2025-04-04 at 04:28:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:28:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "S-II and S-IVB ignition time Saturn V" +2025-04-04 at 04:28:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-04 at 04:28:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo S-II S-IVB ignition time +2025-04-04 at 04:28:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-04 at 04:28:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts