| 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 | 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. |
MSC-02680
CHANGE SHEET
FOR
NASA-MSC INTERNAL REPORT
APOLLO 13 MISSION REPORT
Change 1
# May 1970
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.
# 7.1.6 Batteries
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.
Figure 7.l-l.- Entry battery energy.
# 7.2 LUNAR MODULE
Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of
# 7.1.3 Cryogenic Fluids
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。
| 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 | 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 |
PREPARED BY
# Mission Evaluation Team
APPROVED BY
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970
# TABLE OF CONTENTS
# Section Page
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
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 DESCRIPTIONS·········· A-1
A.1 COMMAND AND SERVICE MODULES .···· A-1
A.2 LUNAR MODULE ······· A-1
A.3 EXPERIMENT EQUIPMENT · A-2
A.4 LAUNCH VEHICLE ······ A-5
A.5 MASS PROPERTIES .. A-5
Section Page
APPENDIX B - SPACECRAFT HISTORIES B-1
APPENDIX C - POSTFLIGHT TESTING C-1
APPENDIX D - DATA AVAILABILITY D-1
APPENDIX E - MISSION REPORT SUPPLEMENTS E-1
REFEREN CES R-1
# 1.0 SUMMARY
The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed.
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 data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory.
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.
The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。
# 2.0 INTRODUCTION
Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material.
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.
In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles.
The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase.
Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I.
Figure 3-l.- Apollo l3 mission profile.
TABLE 3-I.- SEQUENCE OF EVENTS
| 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 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 |
TABLE 4-II.- TRAJECTORY PARAMETERS
| 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 | 159.70E 159.56E | 188 371.38 188 393.19 | 3 065.8 3 093.2 | 79.364 79.934 | 115.464 116.54 |
| Transearth phase Transearth injection |
| Ignition Cutoff Thirdmidcourse correction | Moon Moon | 79:27 :39.0 79:32:02.8 | 3.73N 3.62N | 65.46E 64.77E | 5 465.26 5 658.68 | 4 547.7 5 020.2 | 72.645 64.784 | -116.308 -117.886 |
| Ignition | Earth Earth | 105:18:28.0 105:18:42.0 | 19.63N 19.50N | 136.84W 136.90W | 152 224.32 152 215.52 | 4 457.8 4456.6 | -79.673 -79.765 | 114.134 114.242 |
| Fourthmidcourse correction Ignition Cutofr | Earth Earth | 137:39:51.5 137:40:13.0 | 11.35N 11.34N | 113.39E 113.32E | 37 806.58 37 776.05 | 10 109.1 10 114.6 | -72.369 -72.373 | 116.663 118.660 |
| Service module separation | Earth | 138:01:48.0 | 10.88N | 108.77E | 35 694.93 | 10405.9 | -71.941 | 118.824 |
| Undocking | Earth | 141:30:00.2 | 1.23S | 77.55E | 11 257.48 | 1.7 465.9 | -60.548 | 120.621 |
| Entry interface | Earth | 142:40:45.7 | 28.23S | 173.44E | 65.83 | 36 210.6 | -6.269 | 77.210 |
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.
The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours .
At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours .
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.
# (a) Trans lunar
| 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.
The crew reported a pitch-up disturbance torque was exerted on the command module soon after undocking until the beginning of entry. Most of this time, only low-bit-rate telemetry was available and therefore a detailed analysis is impossible. A 20-minute segment of high-bit-rate data was received just prior to entry, and an unaccountable pitch-up torque of 0.00l deg/sec2 was observed. The possible contributing causes for this torque could have been gravity gradients, atmospheric trimming, venting through the umbilical, venting through the tunnel hatch, and a gradual propellant leak. However, none of these is considered to have been a single cause, and either a combination of these causes was present Or some undetermined venting took place.
Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles.
TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY
| 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 | +21.0 | +3.0 | | |
updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29
coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent.
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 .
| 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.
5.7 REACTION CONTROL
| 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 | | | |
| | -1.2 | 40.4 | |
| | | | +0.4 |
| | | | |
| | | | |
| | | | |
| | | | |
| | | | |
| | | | |
| | | | |
| | | | |
| | | | |
| | | | |
| | | | |
| | | | |
#Earth-centered inertial coordinates. Bystem. \*\*Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance
The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors.
# 6.4.4 Inertial Measurement Unit
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 |
# 6.4.5 Abort Guidance System Performance
Abort guidance system performance was nominal. No instrument calibrations or compensation updates were performed. Uncompensated accelerometer biases and gyro drifts remained within normal operating limits even though heater power was removed from the abort sensor assembly for most of the flight to conserve electrical power. At times, the sensor package temperature was as low as 37? F.
Accelerometer bias shifts associated with the 30-day and 3-day requirements were well within specification. Table 6.4-II contains preflight calibration histories for the initial components of the abort gui dance system.
TABLE 6.4-II.- ABORT GUIDANCE SYSTEM PREINSTALLATION CALIBRATION DATA
| Landing area | Supporta | Remarks |
| Number | Unit |
| Launch site | 1 | LCU | Landing craft utility (landing craft with command |
| Launch abort | 1 | HH-3E | module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida |
| 2 | HH-53C | Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, |
| 1 | ATF | Florida |
| 2 | SH-3 | Helicopters staged from Norfolk NAS, Virginia |
| 1 | DD | USS New |
| 3 | HC-130H | Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, |
| | ’Azores |
| Earth orbit Primary end-of-mission, | 2 2 | DD HC-130H | USS New Fixed wing aircraft staged from Ascension |
| | |
| Mid-Pacific earth | 1 1 | HdT DD | USS Iwo Jima |
| 8 | SH-3D | USS Benjamin Stoddert |
| orbital, and deep- | 2 | | Helicopters staged from USS Iwo Jima |
| space secondary | | HC-130H | Fixed wing aircraft staged from Hickam AFB, Hawaii |
&arotal ship support $=5$ Total aircraft support $\approx$ 23
# 10.3.l Command Module Location and Retrieval
The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards.
The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module.
Figure l0.3-l.- Recovery support at earth landing.
The first reported electronic contact by the recovery forces was through S-band contact by Samoa Rescue 4. A visual sighting report by the Recovery helicopter was received and was followed shortly thereafter by aquisition of the recovery beacon signal by the Recovery, Photo, and Swim l helicopters. Fuel dump was noted and voice contact was made with the descending spacecraft, although no latitude and longitude data were received. The command module landed at 1807 G.m.t. and remained in the stable l flotation attitude. The flashing light was operating and the infiation of the uprighting system commenced about l0 minutes subsequent to landing.
After confirrning the integrity of the command module and the status of the crew, the Recovery helicopter crew attempted to recover the main parachutes with grappling hooks and flotation gear prior to their sinking. Swim l and Swim 2 helicopters arrived on scene and immediately proceeded with retrieval. Swim 2 deployed swimmers to provide flotation to the spacecraft, and Swim l deployed swimmers to retrieve the apex cover, which was located upwind of the spacecraft. The flight crew was onboard the recovery helicopter 7 minutes after they had egressed the command module, and they arrived aboard Iwo Jima at 1853 G.m.t.
Command module retrieval took place at 2l degrees 39.l minutes south latitude and 165 degrees 20.9 minutes west longitude at 1936 G.m.t. One main parachute and the apex cover were retrieved by small boat and brought aboard.
The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston.
Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft.
The following is a chronological listing of events during the recovery operations.
| Event | Time, G.m.t. |
| Apri1 17, 1970 |
| S-band contact by Samoa Rescue 4 Visual contact by Swim 2 | 1801 1802 |
| helicopters | |
| Voice contact by Recovery helicopter | 1803 |
| Visual contact by Relay/Recovery helicopters/ | 1803 |
| Iwo Jima Command module landed, remained in stable I | |
| Swimmers deployed to retrieve main parachutes | 1807 1809 |
| First swimmer deployed to command module | 1816 |
| Flotation collar inflated | 1824 |
| Life preserver unit delivered to lead swimmer | 1831 |
| Command module hatch opened | 1832 |
| Helicopter pickup of flight crew completed | 1842 |
| Recovery helicopter on board Iwo Jima | 1853 |
| Command module secured aboard Iwo Jima | 1936 |
| April 18 |
| Flight crew departed Iwo Jima | 1820 April 20 |
| Flight crew arrival in Houston | 0330 |
| Iwo Jima arrival in Hawaii | April 24 1930 |
| Safing of command module pyrotechnics completed | April_25 0235 |
| Deactivation of the fuel and oxidizer completed | April 26 1928 |
# 10.3.2 Postrecovery Inspection
Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded.
The following discrepancies were noted during the postrecovery inspection:
a. Some of the radioluminescent disks were broken. b. The apex cover was broken on the extravehicular handle side. c. The docking ring was burned and broken. d. The right--hand roll thruster was blistered. e. A yellowish/tan film existed on the outside of the hatch window, left and right rendezvous windows, and the right-hand window. f. The interior surfaces of the command module were very damp and cold, assumed to be condensation; there was no pooling of water on the floor. . Water samples could not be taken from the spacecraft tanks (discussed in section 5.8). h. The postlanding ventilation exhaust valve was open and the inlet valve was closed; the postlanding ventilation valve unlock handle was apparently jammed between the lock and unlock positions (section 14.l.2). i. There was more and deeper heat streaking in the area of the compression and shear pads than has been normally observed.
# 11.0 EXPERIMENTS
# 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA
As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3.
# 11.1.1 Electric Field Measurements
As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges.
Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value.
At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by
Figure ll.l-l.- Field meter location in the laumch site area.
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.
Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately 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 13 launch.
Figure 1l.l-3.- Concluded
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.
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.
Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects .
Figure ll.l-4.- Electrical charge characteristics.
The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V.
# ll.l.2 Very-Low and Low-Frequency Radio Noise
To monitor the low-frequency radio noise, a broad-band antenna system was used at site 7 to feed five receivers, tuned respectively to 1.5 kHz, 6 kHz, 27 kHz, 51 kHz, and 120 kHz.
During launch, a sudden onset of radio noise was observed almost coincidently with the start of the electric field perturbation. This onset was very well marked on all but the 1.5 kHz channel. Following onset, the noise levels at l20 and at 5l kHz tended to decrease slowly in intensity for some 20 seconds. However, the noise levels at 27 and at 6 kHz increased and reached their maxima after about 15 seconds. Furthermore, substantial noise at 1.5 kHz was first apparent at 5 seconds after lift-off and also peaked out in about l5 secorids.
If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower.
# 1l.l.3 Measurement of Pelluric Current
The experiment to measure telluric current consisted of an electrode placed close to the launch site and two electrodes spaced approximately 2500 feet from the base electrode at a 90-degree included angle (shown in figure ll.l-2). The telluric current system failed to detect any launch effects. It was expected that the current would show an increase until the vehicle exhaust plume broke effective electrical contact with ground. The high density of metallic conductors in the ground near the launch site may have functioned as a short circuit, which would have negated the detection of any changes in the current level.
# 11.1.4 Measurement of the Air/Earth Current Density
Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an
altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood.
11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES
The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6).
To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, Advanced Technology Satellite I was out of operation on the day of photography.
TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY
| Description | Completed |
| B C | Television coverage | No |
| Contingency sample collection Selected sample collection | No No |
| D | Evaluation of landing accuracy techniques | No |
| F | | No |
| G | Photographs of candidate exploration sites | |
| H | Extravehicular communication performance | No No |
| I | Lunar soil mechnics | No |
| J | Dim light photography | |
| K | Selenodetic reference point update | No |
| CSM orbital. science photography | No |
| L | Transearth lunar photography | No |
| M | EMU water consumption measurement | No |
| N | Thermal coating degradation | No |
| ALSEPIII | Apollo lunar surface experiments package | No |
| S-059 | Lunar field geology | No |
| S-080 | Solar wind composition | No |
| S-164 | S-band transponder exercise | No |
| S-170 | Downlink bistatic radar observations of the Moon | No |
| S-178 | Gegenschein from lunar orbit | No |
| S-184 | | |
| Lunar surface close-up photography | No |
| T-029 | Pilot describing function | Yes |
8. Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer.
b. Postflight determination of the actual time and location of S-IVF impact to within. l second.
Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3.
Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows:
a. Lunar field geology (S-059)
b. Pilot describing function (T-029)
c. Solar wind composition (S-080)
d. S-band transponder exercise (S-164)
e. Downlink bistatic radar observations of the moon (S-170)
f. Gegenschein observation from lunar orbit (S-178)
g。 Lunar surface closeup photography (S-184)
The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted.
The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3.
The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target (section ll.3).
Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of ±225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an accumulator in the center-engine liquid oxygen line is being incorporated on future vehicles to decouple the line from the crossbeam, and therefore suppress any vibration amplitudes. Addition of a vibration detection system which would monitor structural response in the l4-to-20 Hz range and initiate engine cutoff if vibrations approach a dangerous level. is also under investigation as a backup.
The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted.
# 14.0 ANOMALY SUMMARY
This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission.
# 14.1 COMMAND AND SERVICE MODULES
# 14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure
At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system:
a。 Helium l on quads B and D
b。 Helium 2 on quad D
C. Secondary propellant valves on quads A and C.
Approximately 2-l/2 minutes after the noise, fuel cells l and 3 ceased generating electrical power.
The first indication of a problem in cryogenic oxygen tank 2 occurred when the quantity gage went to a full-scale reading at 46 hours 40 minutes. For the next 9 hours, system operation was nornal. The next abnormal indication occurred when the fans in cryogenic oxygen tank 2 were turned on at 55:53:20. Approximately 2 seconds after energizing the fan circuit, a short was indicated by the current trace from fuel ceil 3, which was supplying power to the oxygen tank 2 fans. Within several additional seconds, two other shorted conditions occurred.
Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers.
The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as the high-gain antenna switched from narrow beam to wide beam, because the panel, when separating, struck and damaged one of the antenna dishes.
Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time.
Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected.
Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l.
The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen
supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems.
A more thorough discussion of this anomaly is presented in reference l.
This anomaly is closed.
# 14.1.2 Postlanding Vent Valve Malfunction
During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen.
The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out.
The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances.
A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results:
a. With the handle extended only l/4 inch or less from the valve
locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve
locked position, the exhaust valve opened but the inlet valve remained
closed.' This condition duplicates that of the position of the handle and
the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from
the valve-locked position, both the inlet and and exhaust valves opened.
Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin.
NASA-S-70-5841
Figure l4-l.- Post-landing vent valve lock.
The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle.
The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found.
To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan.
This anomaly is closed.
# 14.l.3 Shaft Fluctuations in the Zero Optics Mode
Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission.
A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode.
Figure l4-2.- Zero optics mode circuitry.
An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure.
The recurrence of the problem under almost identical circumstances during Apollo l3 indicates that the cause is more likely generic than random and that it is time or vacuum dependent. The susceptibility of the shaft axis rather than the trunnion axis also tends to absolve components common to both axes, such as. the electronics and the motor drive amplifier. The shaft loop has been shown to be more sensitive than the trunnion to harmonics of the 800-hertz reference voltages introduced into the forward loop; however, because the level of the required null offset voltage is well above that available by induction, this mechanism is considered unlikely.
The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation.
Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection.
Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal
14-7
Figure l4-3.- Details of half speed resolver.
Figure l4-4.- One-half speed resolver.
vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on.
The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics.
Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation.
This anomaly is closed.
# 14.l.4 High-Gain Antenna Acquisition Problem
Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station.
When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10.
Starting at 55:00:10 and continuing to 55:00:40, deep repetitive transients approximately every 5 seconds were noted on the phase modulated downlink carrier (fig. l4-5). This type of signature can be caused by a malfunction which would shift the scan-limit and scan-limit-warning function lines, as illustrated in figure 14-5. These function lines would have to shift such that they are both positioned between the antenna manual settings and the true line of sight to earth. Also, the antenna would have to be operating in the auto-reacquisition mode to provide these signatures. The antenna functions which caused the cyclic inflight RF signatures resulting from a shift in the function lines can be explained with the aid of figures 14-5 and 14-6, with the letters A, B, C, and D corresponding to events during the cycle. Starting at approximately 55:00:l0, the antenna was switched from manual to auto reacquisition with the beamwiath switch in the medium-beam position. From point A to the Scan limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual
Figure 14-5.- Shift in scan-limit, scan-limit-warning illustrated.
mode and drive back toward the manual settings until the scan limit warning function line at point C was reached, thereby maintaining wide-beam operation. When the antenna reaches the scan limit warning function line, the system would automatically switch to the medium-beam mode and continue to drive in the manual mode until the manual setting error was nulled out at point A. The antenna would then switch to the auto-track mode and repeat the cycle. The most important feature of this cycle is that the antenna moves at the manual scan rate between points B and D, which is confirmed by the rapid changes in the downlink signal strength.
System testing with a similar antenna and electronics box showed RF signatures comparable to those observed in flight. This consistency was accomplished by placing the target inside the scan limits and the manual setting outside the scan limits. These two positions were separated approximately 35 degrees, which matched the actual angular separation experienced. Under these conditions, the antenna cycled between the target and the manual setting while operating in the auto-reacquisition mode and produced the cyclic RF signature. Since the inflight loss of signal to earth was not near the scan limit, the failure mechanism would be a shift in the Scan-limit function line.
Elements in the scan-limit and scan-limit-warning circuit were shorted and opened to determine the effect on the scan-limit shift.The results of this test shifted the scan-limit functions but did not produce the necessary change in the scan-limit slope. Consequently, a failure in the electronic box is ruled out.
Figure 14-6.- Recorded signal strengths during high-gain antenna operation.
The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees.
The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the
electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees.
The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition.
A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator.
An anomaly report will be published when the analysis is complete.
This anomaly is open.
14.1.5 Entry Monitor System 0.05g Light Malfunction
The entry monitor system 0.05g.light did not illuminate within 3 seconds after an 0.05g condition was sensed by the guidance system. The crew started the system manually as prescribed by switching to the backupposition.
The entry monitor system is designed to start automatically when 0.05g is sensed by the system accelerometer. When this sensing occurs, the 0.05g light should come on, the scroll should begin to drive, and the lrange-to-go counter should begin to count down. The crew reported the light failure but were unable to verify whether or not the scroll or counter responded before the switch was manually changed to the backup mode.
The failure had to be in the light, in the 0.05g sensing mechanism, or in the mode switch, mode switching could also have been premature.
An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive.
A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests.
Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated.
The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry.
Based on these findings, a change is not warranted to existing procedures or hardware on future flights.
This anomaly is closed.
# 14.1.6 Gas Leak in Apex Cover Jettison System
During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units.
The possible causes of the gas leakage include:
a. Out of tolerance parts - Measurement of the failed parts indi
cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was
successful. c. Gap in backup ring - The installation procedure specifies the
backup ring may be trimmed on assembly to meet installation requirements,
Figure 14-7.- Apex cover Jettison system.
but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem.
Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide
Figure 14-8.- Damage from apex jettison thruster.
NA SA-S-70-5849
Figure l4-9.- Plenum side of breech-plenum interface.
sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition.
This anomaly is closed.
Figure. l4-l0.- Tunnel gusset protection.
14.l.7 Reaction Control Isolation Valve Failure
During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation.
The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer"
Figure l4-ll.- Isolation valve circuit.
action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght.
Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired.
This anomaly is closed.
# 14.l.8 Potable Water Quantity Fluctuations
The potable water quantity measurement fluctuated briefly on two occasions during the mission. At about 23 hours, the reading decreased from 98 to 79 percent for about 5 minutes and then returned to a normal reading of approximately l02 percent. Another fluctuation was noted at about 37 hours, at which time the reading decreased from its upper limit to 83.5 percent. The reading then returned to the upper limit in a period of 7 seconds.
Preflight fluctuations of from 2 to 6 percent near the full level were observed once during the countdown demonstration test, and a possible earlier fluctuation of about 4 percent at the half-full level was noted during the flight readiness test.
This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found.
The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement.
This anoma.ly is closed.
# 14.l.9 Suit Pressure Transducer Failure
During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2).
(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure.
During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers.
(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded.
The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing.
This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles.
The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance
assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder.
The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement.
Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation.
To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed.
For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering.
This anomaly is closed.
14.l.l0 Gas Leak in Electrical Circuit Interrupter
During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3.
The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing.
The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include:
a. Sliding friction of the many electrical contact pins, the
several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base
plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres
sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to
the two cartridges e. Physical properties of the attenuator block.
Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary.
This anomaly is closed.
# 14.2 LUNAR MODULE
# 14.2.l Abnormal Supercritical Helium Pressure Rise
During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other
helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively.
The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank.
The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard.
The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process.
The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant.
Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure.
A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity.
This anomaly is closed.
# 14.2.2 Abnormal Descent Stage Noise
At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down.
Figure 14-l5.- Descent stage battery location.
The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition.
The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section.
Evidence indicates that battery 2 may have experienced an electrical fault of some type. The most probable condition is electrolyte leaking from one or more cells and bridging the high-voltage or low-voltage terminal to the battery case (fig. 14-17). This bridging results in water electrolysis and subsequent ignition of the hydrogen and oxygen so generated. The accompanying "explosion" would then blow off or rupture the seal of the battery lid and cause both a thump and venting of the free liquids in the battery case, resulting in "snowflakes."
Postflight tests have shown the following:
a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes.
For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage.
NASA-S-70-5859
Figure 14-l7.- Descent battery terminal configuration.
The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing.
In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible.
The designs of other Apollo batteries have been reevaluated, and all are considered safe except the lunar module ascent batteries and the lunar surface drill battery. The ascent batteries and a new battery to be installed in subsequent service modules will receive the Same corrective action applied to the descent battery. The lunar surface drill battery, which previously was unpotted, will be. potted.
This anomaly is closed.
# 14.2.3 Descent Battery 2 Malfunction Light On
The battery malfunction light illuminated at about l00 hours with a corresponding master alarm. The malfunction, isolated to battery 2, could have been caused by an overcurrent, a reverse-current condition, an overtemperature condition, or possibly an erroneous indication. The logic for these malfunction conditions is shown in figure l4-l8.
NASA-S-70-5860
Figure l4-l8.- Battery 2 malfunction circuit.
A battery overcurrent can be ruled out because automatic removal of the battery from the bus would have occurred.
A reverse-current condition can be ruled out because, if the battery is removed from and reapplied to the bus, the reverse-current circuit has a built-in delay of about 5 seconds before the reverse-current relay is again activated to illuminate the light. Battery power was removed from and replaced on the bus in flight, and the light immediately illuminated again when the battery was reconnected.
An over-temperature condition can be ruled out because, after the battery was replaced on the bus, the light remained illuminated for a brief period and then began flickering intermittently. A flickering light cannot be caused by the temperature sensing switch because of a temperature hysteresis of approximately $_{20}\circ$ F in the switch. The water glycol loop temperature also indicated that the battery temperature was normal.
Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here.
Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required.
This anomaly is closed.
# 14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak
During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction.
The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring
Figure 14-l9.- Oxygen-supply system.
(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination.
The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly.
Figure l4-20.- Ascent stage tank shutoff valve:.
Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves.
This anomaly is closed.
# 14.2.5 Cracked Window Shade
The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade.
Figure 14-21.- Cracked left-hand window shade.
Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight.
The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight.
This anomaly is closed.
# 14.3 GOVERNMENT FURNISHED EQUIPMENT
14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera
For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies.
To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly.
This anomaly is closed.
# 14.3.2 Failure of the Interval Timer Set Knob
The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin.
This anomaly is closed.
# 14.3.3 Improper Nasal Spray Operation
When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays.
Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops.
This anomaly is closed.
# 15.0 CONCLUS IONS
The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report.
a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank.
b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles).
c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime.
d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency.
e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived.
The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented.
# A.1 COMMAND AND SERVICE MODULES
The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged urine trans. fer line. Also included was a lunar topographic camera, which could be installed in the command module hatch window for high resolution photography of the lunar surface from orbit. The camera provided a 4.5-inch film format and had an 18-inch focal length and image-motion compensation. The photographs would yield a resolution of approximately l2 feet and would include a l5-mile square area on the surface for each frame exposed.
# A.2 LUNAR MODULE
The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag.
The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage.
# A.3 EXPERIMENT EQUIPMENT
The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13.
The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay.
NASA-S-70-5864
Figure A-l.- Experiment subpackage number l.
# A.3.1 Heat Flow Experiment
The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements.
The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring.
Figure A-2.- Experiment subpackage number 2.
A.3.2 Charged Particle Lunar Environment Experiment
The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds.
# A.3.3 Cold Cathode Gage Experiment
The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms.
The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover.
# A.4 LAUNCH VEHICLE
Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation.
# A.5 MASS PROPERTIES
Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated.
# TABLE A-I.- MASS PROPERTIES
| Event | Weight, 1b | Center of gravity, in. | Moment or inertia, slug-ft2 | Product of inertia, slug-ft2 |
| X | | Z | | | | IxY | | |
| Lift-off | 110 252.4 | 847.4 | 2.4 | 3.7 | 67646 | 1 175 539 | 1 178 016 | 2906 | 8047 | 3711 |
| Earth orbit insertion | 101 261.2 | 807.4 | 2.6 | 4.1 | 66770 | 718 686 | 721 213 | 5157 | 11945 | 3688 |
| Command&servicemodules Lwnar module | 63 720.3 33499.1 | 934.5 1237.0 | 4.0 -0.1 | 6.5 0.0 | 33995 22457 | 76486 24654 | 79123 25255 | ~1746 | -126 95 | 3221 235 |
| Totaldocked | 97 219.4 | 1038.7 | 2.6 | 4.3 | 56 736 | 534890 | 538009 | -8142 | -9376 | 3585 |
| First midcourse correction Ignition Cutoff | 97 081.5 96 851.1 | 1038.9 1039.0 | 2.6 2.6 | 4.2 4.2 | 56 629 56 508 | 534493 534 139 | 537 635 537 380 | -8192 -8189 | -9305 -9282 | 3620 3587 |
| Cryotenic oxygen tank incitent Before | 96 646.9 | 1039.2 | 2.6 | 4.2 | 56 321 | 533499 | 536 766 | -8239 | -9244 | 3636 |
| After Second midcourse correction | 96 038.7 | 1040.7 | 3.0 | 3.9 | 57248 | 533 927 | 537 251 | -8269 | 669- | -3709 |
| Ignition Cutoff | 95 959.9 95 647.1 | 378.8 379.4 | 4.9 5.0 | 0.7 0.7 | 57205 57006 | 516443 513919 | 521 180 518700 | 11617 11553 | 2659 2651 | 3286 3285 |
| Transearth injection Ignition Cutoff | 95 424.0 87456.0 | 379.7 398.4 | 5.0 5.5 | 0.7 0.8 | 56 866 51778 | 512 837 431285 | 517 560 437119 | 11370 9443 | 2495 2222 | 3255 3249 |
| Thirdmidcourse correction Ignition Cutoff | 87 325.3 87 263.3 | 398.7 398.9 | 5.5 5.5. | 0.8 0.8 | 51 681 51642 | 430 123 429353 | 435930 435 169 | 9244 9227 | 2048 | 3215 |
| Fowrth midcourse correction Ignition | 87 132.1 | 399.1 | 5.5 | 0.8 | 51 553 | 428 322 | 434105 | 9069 | 2045 1911 | 3215 3191 |
| Cutoff Command&servicemodule | 87 101.5 | 399.2 | 5.6 | 0.8 | 51538 | 428219 | 433990 | 9065 | 1910 | 3192 |
| b separation Before | 87 057.3 | 399.3 | 5.6 | 0.8 | 51 517 | 428065 | 433 819 | 9058 | 1909 | 3194 |
| After (command module/ lunar module) | 37 109.7 | 251.5 | 2.2 | -0.3 | 24048 | 92418 | 93.809 | 2362 | 686= | 9 |
| Commandmodule/lunar | | | | | | | | | | |
| module separation b | 37 014.6 | | | | | | | | | |
| Before | 12 367.6 | 252.9 1039.9 | 1.9 | -0.6 | 23926 | 93993 | 95514 | 2188 | -963 | -35 |
| After (command module) | | | 0.3 | 6.1 | 581.5 | 5 258 | 4636 | 31 | 409 | 20 |
| Entry | 12 361.4 | 1039.9 | 0.3 | 6.0 | 5 812 | 5254 | 4635 | 31 | -407 | 21 |
| Drogue deployment | 11 869.4 | 1038.7 | 0.3 | 6.0 | 5727 | 5002 | 4405 | 33 | -382 | 24 |
| Main parachute deployment | 11 579.8 | 1038.6 | 0.5 | 5.3 | 5590 | 4 812 | 4346 | 27 | -319 | 41 |
| Landing | 11 132.9 | 1036.6 | 0.5 | 5.2 | 5526 | 4531 | 4046 | 25 | -328 | 42 |
ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases.
The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2.
The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4.
Figure B-l.- Checkout flow for command and service modules at contractor's facility.
NASA-S-70-5867
Figure B-2.- Command and service module checkout history at Kennedy Space Center.
NASA-S-70-5868
Figure B-3.- Checkout flow for lunar module at contractor's facility.
Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center.
The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included.
# TABLE C-I.- POSTFLIGHT TESTING SUMMARY
| ASHUR | Purpose | Tests performed | Results |
| Environmental Control |
| 109007 | To determine contaninates present or damage incurred in 9o0 psi system | Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator | |
| 109008 | To determine contaninates present in residual oxygen in surge tank snd repressurization package | Withdraw sample and analyze for contaminates | No rigrificant difference from the araiysis per- formed at:adine |
| 109016 | To investigate the failure of the postlanding ventilation valve to cycle open | Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis | Not complete |
| 109020 | Todetermine the cause of failure othe suit pressure transaucer | Perform calibration check,dis- assembly,and failure anaysis | Not compiete |
| 109021 | Todetermine the cause of failure o!the potable water transducer | Remove,disassemble,and per- form failure analysis | Hot complete |
| 109015 | To investigate the cause for optics coupling display unit indications ofoptics movement during the | Guidance and Navigation Performance check of zero optics mode operation | Unable to perform tests on optical unit due to sslt water contanination |
| 109018 | zero optics mode To investigate the failure ofthe 0.0)g indication during entry | Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring | Not complete |
Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted.
TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY
| Supplement number | Title | Publi cation date/status |
| Apollo 7 |
| 1 2 | Trajectory Reconstruction and Analysis Communication System Performance | May 1969 June 1969 |
| 3 | Guidance, Navigation, and Control System Performance Analysis | November 1969 |
| 4 5 | Reaction Control System Performance Cancelled | August 1969 |
| 6 | Entry Postflight Analysis | December 1969 |
| Apollo 8 |
| 1 | Trajectory Reconstruction and Analysis | December 1969 |
| 2 | Guidance, Navigation, and Control System Performance Analysis | November 1969 |
| 3 | Performance of Command and Service Module | March 1970 |
| 4 | Reaction Control System Service Propulsion System Final Flight | September 1970 |
| 6 | Evaluation Analysis of Apollo 8 Photography and | December 1969 |
| 7 | Visual Observations Entry Postflight Analysis | December 1969 |
| Apollo 9 |
| 1 | Trajectory Reconstruction and Analysis | November 1969 |
| 2 | Command and Service Module Guidance, Navi- gation, and Control System Performance | November 1969 |
| 3 | Lunar Module Abort Guidance System Perform- ance Analysis | November 1969 |
| 4 | Performance of Command and Service Module Reaction Control System | Apri1 1970 |
| 5 | Service Propulsion System Final Flight Evaluation | December 1969 |
| 6 | Performance of Lunar Module Reaction Control System | Final review |
| 7 | Ascent Propulsion System Final Flight Evaluation | December 1969 |
| 8 | Descent Propulsion System Final Flight Evaluation | September 1970 |
| 9 | Cancelled | |
| 10 | Stroking Test Analysis | December 1969 |
| 11 | Communications System Performance | December 1969 |
| 12 | Entry Postflight Analysis | December 1969 |
| Supplement number | Title | Publication date/status |
| Apollo 10 |
| 1 | Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System | March 1970 |
| 2 | Performance Analysis | December 1969 |
| 3 | Perfornance of Command and Service Module Reaction Control System | Final review |
| 7 | Service Propulsion System Final Flight | September 1970 |
| 5 | Evaluati on Performance of Lunar Module Reaction Control | Final review |
| 6 | System Ascent Propulsion System Final Flight | January 1970 |
| 7 | Evaluation Descent Propulsion System Final Flight | January 1970 |
| 8 | Evaluati on Cancelled | |
| 9 | Analysis of Apollo lo Photography and Visual Observations | In publication |
| 10 11 | Entry Postflight Analysis Communi cations System Performance | December 1969 December 1969 |
| Apollo ）11 |
| 2 3 4 | Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation | May 1970 September 1970 Review Review |
| Supplement number | Title | Publication date/status |
| Apollo 12 Trajectory Reconstruction and Analysis |
| 1 2 3 | Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight | September 1970 September 1970 |
| 4 | Evaluation Ascent Propulsion System Final Flight Evaluation | Preparation Preparation |
| 5 6 | Descent Propulsion System Final Flight Evaluation | Preparation |
| 7 | Apollo l2 Preliminary Science Report Landing Site Selection Processes | July 1970 Final review |
| Apollo 13 | |
| 1 | Guidance, Navigation, and Control System Performance Analysis | Review |
| 2 | Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis | Preparation |
# REFERENCES
1. Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970.
2. Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970.
3. Marshall Space Flight Center, Kennedy Space Center, Manned Spacecraft Center: Analysis of Apollo l2 Lightning Incident, MSC-01540. February 1970.
4. ICSU/IUGG Committee on Atmospheric Sciences: Report of the Study Conference on the Global Atmospheric Research Program, 1967.
5. Bulletin of the American Meteorological Society, Vol. 50, No. 7: Cloud Height Contouring from Apollo 6 Photography, by V. S. Whitehead, I. D. Browne, and J. G. Garcia. 1969.
6. Defense Supply Agency, Washington, D. C.: Military Standardization Handbook_ Optical Design, MIL HDBK-14l. 1962.
7. NASA Headquarters: Apollo Flight Mission Assignments. OMSF M-D MA500-11 (SE 010-000-1). 0ctober 1969.
8. Manned Spacecraft Center: Mission Requirement, H-2 Type Mission (Lunar Landing). SPD9-R-053. November 10, 1969.
# APOLLO SPACECRAFT FLIGHT HISTORY
(Continued from inside front cover)
| Mi ssion. Apollo 4 | Spacecraft | Description | Launch date | Launch site |
| SC-017 LTA-10R | Supercircular entry at lunar. return velocity | Nov. 9, 1967 | Kennedy Space Center, Fla. |
| Apollo 5 | LM-1 | First lunar module flight | Jan. 22, 1968 | Cape Kennedy, Fl&. |
| Apollo 6 | SC-020 LTA-2R | Verification of closed-loop emergency detection system | Apri1 4, 1968 | Kennedy Space Center, Fia. |
| Apo1lo7 | CSM 101 | First manned flight; earth-orbital | 0ct.11，1968 | Cape Kennedy, Fl&. |
| Apollo 8 | CSM 103 | First manned lumar orbital flight; first manned Saturn V launch | Dec.21,1968 | Kennedy Space |
| Apollo 9 | CSM 104 LM-3 | First manned lunar module flight; earth orbit rendezvous; EVA | Mar.3,1969 | Kennedy Space Center,Fla. |
| Apollo 10 | CSM 106 t-WT | First lumar orbit rendezvoua; low pass over lunar surface | May 18, 1969 | Kennedy Space Center, Fla. |
| Apollo 11 | CSM 107 LM-5 | First lunar landing | July 16, 1969 | Kennedy Space |
| Apollo 12 | CSM 108 LM-6 | Second lunar landing | Nov. 14, 1969 | Center, Fla. Kennedy Space Center. F18 |
POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION
NASA-Manned Spacecraft Center Houston, Texas 77058
ATTN: PT2(office Symbol)
(Continued from inside front cover)
| Mi ssion | Spacecraft | Description | Launch date | Launch site |
| Apollo4 | SC-017 LTA-10R | Supercircular entry at lunar | Nov.9,1967 | Kennedy Space Center, Fla. |
| Apollo 5 | LM-1 | return velocity First lunar module flight | Jan.22,1968 | Cape Kennedy, |
| Apollo 6 | SC-020 LTA-2R | Verification of closed-loop | April 4, 1968 | Fla. Kennedy Space Center, Fla. |
| Apollo7 | CSM 101 | emergency detection system First manned flight; | Oct.11，1968 | |
| Apol1o 8 | CSM 103 | earth-orbital First manned lunar | Dec.2l,1968 | Cape Kennedy, Fla. Kennedy Space |
| Apol1o9 | CSM 104 | orbital flight; first manned Saturn V launch First manned lunar | | |
| Apollo 10 | LM-3 | module flight; earth orbit rendezvous; EVA | Mar.3,1969 | Kennedy Space Center, Fla. |
| CSM 106 LM-4 | First lunar orbit rendezvous; low pass over lumar surface | May 18, 1969 | Kennedy Space Center, Fla. |
| Apollo 11 | CSM 107 LM-5 | First lunar landing | July 16, 1969 | Kennedy Space |
| Apollo 12 | CSM 108 | Second lunar landing | | Center, Fla. |
| Apollo 13 | LM-6 | | Nov. 14, 1969 | Kennedy Space Center, Fla. |