Question: On what date did the mission PA-1 experience a first pad abort?

Answer: Nov.7, 1963

Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966


Question: Where did the mission AS-102 take place?

Answer: Cape Kennedy, Fla.

Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966


Question: Where was the AS-201 SC-009 Supercircular entry located on February 26, 1966?

Answer: N. Mex.

White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla.


Question: What type of entry did the AS-202 SC-011 experience on August 25, 1966?

Answer: Supercircular entry with high heat load

White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla.


Question: What is the title of the report?

Answer: APOLLO 13 MISSION REPORT

James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program

After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted."

In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated.

NOTE: A black bar in the margin of affected pages indicates the information that was changed or added.


Question: Who is the Manager, Apollo Spacecraft Program?

Answer: James A. MeDivitt Colonel, USAF

James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program

After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted."

In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated.

NOTE: A black bar in the margin of affected pages indicates the information that was changed or added.


Question: How many ampere-hours of energy remained in the batteries at landing?

Answer: 29

The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained.

Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of

Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。


Question: How long did it take for oxygen tank 1 to be depleted after the incident?

Answer: 2 hours

The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained.

Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of

Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。


Question: What was the estimated total energy transferred to the command module?

Answer: approximately 129 ampere hours

operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking.

Figure 7.2-l.- Lunar module water usage.

Figure 7.2-2.- Lunar module total battery capacity during flight.

Figure ll.l-2.- Field meter locations in the proximity of the launch complex.

gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6.


Question: What was the amount of ampere hours remaining in the lunar module batteries at the time of undocking?

Answer: 410

operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking.

Figure 7.2-l.- Lunar module water usage.

Figure 7.2-2.- Lunar module total battery capacity during flight.

Figure ll.l-2.- Field meter locations in the proximity of the launch complex.

gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6.


Question: What was found on the moving portion of the records from sites 8 and 9?

Answer: anything

Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off.

Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch.

No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations.


Question: At which stations were no significant perturbation in the electric field produced by the launch cloud?

Answer: 4 or 5

Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off.

Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch.

No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations.


Question: What type of signals were recorded at the launch site during the afternoon of launch day?

Answer: Sporadic signals

The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12.

The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood.


Question: What was the weather situation described as before the launch of the Apollo 13 vehicle?

Answer: Marginal

The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12.

The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood.


Question: What is the equilibrium potential that a conventional jet aircraft can approach?

Answer: a million volts

It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more.

Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation


Question: What is the publication status of the document titled "Preparati on 4 Ascent Propulsion System Final Flight Evaluation"?

Answer: Final Flight Evaluation

It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more.

Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation


Question: What was the publication date of the report titled "Trajectory Reconstruction and Analysis"?

Answer: March 1970

Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5


Question: What was the status of the report titled "Analysis of Apollo 10 Photography and Visual In"?

Answer: Cancelled

Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5


Question: What was the subject of the document titled "Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation"?

Answer: Ascent Propulsion System

Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970


Question: In which month was the Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis conducted?

Answer: December

Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970


Question: What year was the document approved by the Mission Evaluation Team?

Answer: 1970

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970


Question: Where is the National Aeronautics and Space Administration Manned Spacecraft Center located?

Answer: Houston, Texas

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970


Question: What is the section that discusses the performance of the command and service module?

Answer: 5.0

1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7


Question: What is the section that describes the training of the pilots?

Answer: 8.1

1.0 SUMMARY 1-1 2.0 INTRODUCTION·. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ··· 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .· 5-1 5.2 ELECTRICAL POWER ···· 5-2 5.3 CRYOGENIC STORAGE.··· 5-3 5.4 COMMUNICATIONS EQUIPMENT · 5-4 5.5 INSTRUMENTATION.······· 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .· 5-5 5.7 REACTION CONTROL.······· 5-11 5.8 ENVIRONMENTAL CONTROL .·. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ··· 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .· 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ··· 6-8 6.7 ENVIRONMENTAL CONTROL.··· 6-9 7.0 MISSION CONSUMABLES ·····. ··、· 7-1 7.1 COMMAND AND SERVICE MODULES .···· 7-1 7.2 LUNAR MODULE ····· 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .·.. 8-7


Question: What is the title of the section that discusses the physical examinations of the astronauts?

Answer: 9.3 PHYSICAL EXAMINATIONS

8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE


Question: What is the page number where the seismic detection of third stage lunar impact experiment is discussed?

Answer: 11-9

8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ····· 8-11 8.10 ENTRY AND LANDING.··. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ······· · 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ···· 10-1 10.2 NETWORK．······· 10-2 10.3 RECOVERY OPERATIONS...·.·. ··· 10-2 11.0 EXPERIMENTS·····．···.····. ·· 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES．·．··．·．··．·．··．·· 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ·· 12-1 13.0 LAUNCH VEHICLE SUMMARY·......·......... 13-1 14.0 ANOMALY SUMMARY ·········· 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ·· 14-1 14.2 LUNAR MODULE ············ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ··· 14-36 15.0 CONCLUSIONS····.···.··．·.．·· ·· 15-1 APPENDIX A - VEHICLE


Question: What was the time of launch in e.s.t.?

Answer: 2:13:00 p.m.

The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological


Question: Who was substituted as the Command Module Pilot two days before launch?

Answer: John L. Swigert, Jr.

The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological


Question: How long did it take for the pressure in cryogenic oxygen tank 2 to rise at an abnormally high rate?

Answer: Approximately 56 hours

At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry.


Question: What was the reason for the immediate abort of the mission?

Answer: Loss of oxygen and primary power in the service module

At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry.


Question: What section of the report contains a discussion of the abort profile?

Answer: Section 3

Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7.

A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified.


Question: Where can the current status of all Apollo mission supplements be found?

Answer: Appendix E

Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7.

A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified.


Question: What was the location of the resultant landing at earth?

Answer: Indian Ocean

After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with


Question: How many hours of support did the lunar module systems intend to provide for the crew after the resultant landing?

Answer: 90 hours

After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with


Question: What was the location of the landing site?

Answer: The South Pacific

The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control.


Question: How many hours before entry was the final midcourse correction performed?

Answer: 5

The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control.


Question: What was the power source for the three entry batteries?

Answer: Lunar module power

The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative


Question: Where did the lunar module, including the radioisotope thermoelectric fuel capsule, impact?

Answer: The open sea between Samoa and New Zealand

The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative


Question: What was the planned orbital altitude of the pericynthion after the first major spacecraft maneuver?

Answer: 60 miles

As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II.

TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS


Question: What was the achieved pericynthion altitude at translunar injection?

Answer: 415.8 miles

As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II.

TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS


Question: What unit is used to measure altitude above the lunar surface?

Answer: feet or miles

Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial


Question: What is the reference system for the space-fixed velocity vector?

Answer: body-centered, inertial reference coordinate system

Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect，feet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial


Question: What is the unit of measurement for the maximum altitude above the oblate earth model?

Answer: mile

velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg


Question: What is the acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane?

Answer: Inclination

velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg


Question: What was the altitude of the spacecraft above the launch pad at the time of S-IVB second cutoff?

Answer: 175.71

Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N


Question: What was the space-fixed flight-path angle at the time of translunar injection?

Answer: 59.318

Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N


Question: What is the time of the Transearth phase?

Answer: 79:27 :39.0


Question: What is the latitude of the Earth at the time of the service module separation?

Answer: 10.88N


Question: At what time was the S-IVB maneuver to achieve lunar impact initiated?

Answer: 6 hours

The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0.


Question: What was the unexplained velocity increase of the S-IVB detected by tracking data at approximately 19 hours 17 minutes?

Answer: 5 ft/sec

The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0.


Question: What was the time at which the first transearth midcourse correction was performed?

Answer: 105:18:28

The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees .

Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12.


Question: What was the velocity change produced by the transearth midcourse correction?

Answer: 7.8 ft/sec

The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees .

Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12.


Question: What was the ignition time of the S-IVB?

Answer: 2:35:46.4

Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57


Question: What was the latitude of the pericynthion arrival?

Answer: 1.47N

Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57


Question: What was the time of the maneuver system ignition?

Answer: 79:27:39

Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46


Question: What was the latitude of the spacecraft at entry arrival time?

Answer: 28.22S

Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46


Question: What was the time of the final midcourse correction maneuver?

Answer: 5 hours before entry

The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel.

The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point.


Question: Where did the spacecraft land?

Answer: The Pacific Ocean

The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel.

The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point.


Question: What type of propulsion system is discussed in this section?

Answer: Service propulsion

The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary.


Question: Where are the details of the anomalies discussed?

Answer: In the Anomaly Summary

The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary.


Question: How many times were batteries A and B charged during the flight?

Answer: Three

Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries.


Question: What was the source of power for main bus A after the cryogenic oxygen incident?

Answer: Battery A

Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries.


Question: What was the issue with the fuel cell flow indicators before lift-off?

Answer: Erratic readings

Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal.

During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l.


Question: Which fuel cell continued to operate after the loss of oxygen pressure in tank 2?

Answer: Fuel cell 2

Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal.

During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l.


Question: What was the average current of the fuel cells during the mission?

Answer: 24 amperes

During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts.


Question: What was the average bus voltage of the fuel cells during the mission?

Answer: 29.4 volts

During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts.


Question: What type of communications were used after translunar injection?

Answer: S-band

Consumable quantities in the cryogenic storage system are discussed in section 7.l.

The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good.

Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4.


Question: How long was the spacecraft powered down?

Answer: approximately 58 hours

Consumable quantities in the cryogenic storage system are discussed in section 7.l.

The communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good.

Prior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4.


Question: What was the time period during which the communications system was powered up to transmit high-bit-rate telemetry data using the omnidirectional antennas?

Answer: From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00

At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data.

From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery.


Question: What was the cause of the damage to the high-gain antenna?

Answer: The loss of a Service module outer panel

At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data.

From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery.


Question: What was the discrepancy in the suit pressure measurement?

Answer: 0.5 psi below cabin pressure

The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement · operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident.


Question: Where is the discussion on the potable water quantity measurement anomaly located?

Answer: section 14.l.8

The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement · operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident.


Question: What was the reason for the failure of the service propulsion auxiliary propellant gaging system?

Answer: Fuel leakage into the point sensor module within the tank.

The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected.


Question: Did the failure of the service propulsion auxiliary propellant gaging system affect the performance of the mission?

Answer: No

The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected.


Question: How long did the initial separation from the S-IvB take?

Answer: 4.28 seconds

All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll.


Question: What was the peak rate disturbance in roll during docking?

Answer: 0.60 deg/sec

All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll.


Question: What was the reason for the unsuccessful passive thermal control mode attempt at 7:43:02?

Answer: The roll rate was established using one rather than two roll engines, as required by the checklist.

The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine


Question: What triggered the unplanned minimum impulse engine firing at 32:21:49?

Answer: The roll manual attitude switch was changed from the rate-command position to the acceleration-command position.

The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine


Question: What was the time at which the passive thermal control mode was attempted?

Answer: 32:2l:49

firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally.

NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode.


Question: What were the adverse effects of the two extraneous firings shown in figure 5.6-1?

Answer: Adverse effects of two extraneous firings

firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally.

NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode.


Question: What caused a computer restart during the oxygen tank incident?

Answer: A voltage transient

At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly .

The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table.


Question: How long did it take for the rate and attitude errors to be reduced to a nulled condition?

Answer: 75 seconds

At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly .

The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table.


Question: What is the time of cutoff for the mission?

Answer: 30 : 40 :53.14

Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1  +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state  Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll

Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed.


Question: What is the initial pitch of the spacecraft?

Answer: 0.95

Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1  +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state  Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll

Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed.


Question: What is the time of the first entry in the table?

Answer: 00:45

preferred alignment DRererence metrix (REFSMAT) CCoarse alignment


Question: What is the star used in the coarse alignment at time 40:43?

Answer: 36Vega,40Altair

preferred alignment DRererence metrix (REFSMAT) CCoarse alignment


Question: What was the difference in velocity between the S-IvB instrument unit and the command module platform during earth ascent?

Answer: 75-ft/sec

Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift


Question: What was the cause of the discrepancy in the Y-axis error magnitude?

Answer: The magnitude of the null bias drift

Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift


Question: What is the X-Scale factor error in ppm?

Answer: 2

Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7


Question: What is the acceleration drift in mERU/g for the input axis in the Z-direction?

Answer: -1.0

Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 Bia，cm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1？ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis，mERU/g..， -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi日,mERU/g +19.17 7.14 7


Question: At what time was power to the guidance and navigation system removed?

Answer: 58 hours

After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments .


Question: What was the new value of the Z-axis accelerometer bias after the update?

Answer: minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$

After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments .


Question: What was the initial accelerometer bias before translunar injection?

Answer: +0.008

Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5

Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table.

Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015


Question: What was the acceleration drift of the input axis?

Answer: 9.0

Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.· Y Z 96- 116 37 116 Lt- 116 Null bias drift， mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5

Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table.

Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015


Question: How much propellant was used for the initial separation from the S-IVB?

Answer: 55 pounds

All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission.


Question: What was the difference between predicted and actual propellant usage before the tank anomaly?

Answer: 33 pounds

All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission.


Question: What was the state of the propellant isolation valves during system decontamination at Hawaii?

Answer: Normal (closed)

System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7.

All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation.


Question: What was used to close the system 1 fuel isolation valve?

Answer: Power from ground servicing equipment

System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7.

All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation.


Question: At what time was the suit compressor turned off?

Answer: 56:19:58

The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power.


Question: What was the pressure in the surge tank when it was isolated?

Answer: 858 psia

The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power.


Question: What was the duration of nominal S-band communications from system actuation?

Answer: Approximately 58 hours

S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality.


Question: What was the primary configuration used for communications?

Answer: Low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband

S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality.


Question: What was the temperature range of the antenna in the passive thermal control mode?

Answer: plus and minus $25^{\circ}$ F

The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。

System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit.


Question: What was the temperature of the antenna when the antenna heaters were turned off?

Answer: minus 66° F

The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。

System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit.


Question: What was the duration of the first period of throttle increase during the transearth injection maneuver?

Answer: 5 seconds

The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs.


Question: What system was used to control yaw during the first transearth midcourse correction?

Answer: abort guidance system attitude-hold mode

The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs.


Question: What was the duration of the third midcourse correction?

Answer: 34.23

Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44         Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg  Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60


Question: What was the maximum rate excursion of the pitch gimbal drive actuator in degrees per second?

Answer: -0.6

Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# 人 +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44         Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 ±0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg  Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60


Question: What was the standard deviation of the scale factor error for the X-axis accelerometer?

Answer: -681

The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table.

Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0


Question: How many samples were used to calculate the average bias for the Y-axis accelerometer?

Answer: 5

The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table.

Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis，mERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.· -5.38 2.37 4 -5.5 -4.0


Question: What is the standard deviation of the accelerometer bias in the X direction?

Answer: 16.3

Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr


Question: What is the mean gyro spin axis mass in the X direction?

Answer: 0.86

Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 人 -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee， deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 人 -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr


Question: What was the initial propellant consumption of the reaction control system?

Answer: 467 pounds

The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds.

About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open.

During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。


Question: What was the duration of the decrease in system-A propellant manifold pressures?

Answer: 4 or 5 seconds

The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds.

About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open.

During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。


Question: What was the duration of the transearth injection maneuver?

Answer: 264 seconds

The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained.


Question: At what throttle position was the pressurization isolation Solenoid closed?

Answer: l2 percent of full thrust

The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained.