apollo 13 mission operation report

8/8/2019 Apollo 13 Mission Operation Report

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_I

TO: A/Administrator

PrelaunchMission Operation ReportNo. M-932-70- 13

31 March 1970

FROM: MA/Apollo Program Director

SUBJECT: Apollo 13 Mission (AS-508)l

On 11 April 1970, we plan to launch Apollo 13 from Pad A of Launch Complex 39 atthe Kennedy Space Center. This will be the third manned lunar landing mission and is. targeted to a preselected point in the Fra Mauro Formation.

Primary objectives of this mission include selenological inspection, survey, andsampling of the ejecta blanket thought to have been deposited during the formation ofthe lmbrium basin; deployment and activation of an Apollo Lunar Surface ExperimentsPackage; continuing the development of mans capability to work in the lunar environ-ment; and obtaining photographs of candidate lunar exploration sites. Photographicrecords will be obtained and the extravehicular activities will be televised.The IO-day mission will be completed with landing in the Pacific Ocean. Recoveryand transport of the crew, spacecraft, and lunar samples to the Lunar Receivingr/- Laboratory at the Manned Spacecraft Center will be conducted under quarantineprocedures that provide for biological isolation.

EL~k-Rocco A. Petrone

~s!Zc~;e%%%f!tor forManned Space Flight


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FOREWORD

MISSION OPERATION REPORTS are published expressly for the use of NASA SeniorManagement, as required by the Administrator in NASA Instruction 6-2-10, dated15 August 1963. The purpose of these reports is to provide NASA Senior Management

with timely, complete, and definitive information on flight mission plans, and toestablish official mission objectives which provide the basis for assessment o f missionaccomplishment.Initial reports are prepared and issued for each flight project just prior to launch.Following launch, updating reports foreach missionare issued to keepGeneral Manage-ment currently informed of definitive mission results as provided in NASA Instruction6-2-10Primary distribution of these reports is intended for personnel having program/projectmanagement responsibilities which sometimes results in a highly technical orientation.The Office of Public Affairs publishes a comprehensive series of pre-launch and post-launch reports on NASA flight missions which are available for dissemination to thePress.

APOLLO MISSION OPERATION REPORTSare published in two volumes: theMISSIONOPERATION REPORT (MOR) ; and the MISSION OPERATION REPORT, APOLLOSUPPLEMENT. This format was designed to provide a mission-oriented document inthe MOR, with supporting equipment and facility description in the MOR, APOLLOSUPPLEMENT. The MOR, APOLLO SUPPLEMENT is a program-oriented referencedocument with a broad technical description of the space vehicle and associated equip-ment, the launch complex, and mission control and support facilities.

Published and Distributed byPROGRAM and SPECIAL REPORTS DIVISION (XP)

EXECUTIVE SECRETARIAT - NASA HEADQUARTERS


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CONTENTSPage

Summary of Apollo/Saturn Missions . . . . . . . . . . . . . . . . . . . . . . 1.NASA OMSF Primary Mission Objectives for Apollo 13. . . . . . . . . . . . 2

. Detailed Objectives and Experiments . . . . . . . . . . . . . . . . . . . . . 3Launch Countdown and Turnaround Capability, AS-508. . . . . . . . . . . . 5

Flight Mission Description. . . . . . . . . . . . . . . . . . . . . . . . . . . 6

Contingency Operations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29Mission Support . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32

Configuration Differences. . . . . . . . . . . . . . . . . . . . . . . . . . . 39I-Flight Crew . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41

Mission Management Responsibility . . . . . . . . . . . . . . . . . . . . . . 45

Abbreviations and Acronyms. . . . . . . . . . . . . . . . . . . . . . . . . . 46

-.

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Figure No.1

. 2

345678,--

91011 LM Ascent Through Docking1213

1415161718

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LIST OF FIGURESTitle

Apollo 13 Flight ProfileLunar Orbit Insertion and Descent Orbit Insertion

GeometryLM Descent Orbital EventsApollo 13 Lunar Surface Activities SummaryApollo 13 EVA-l Time1 ineDeployed Modularized Equipment Stowage AssemblyDeployed S-band AntennaApollo 13 ALSEP Radioisotope Thermoelectric

Generator (Unfueled)EVA- 1 Equipment DeploymentApollo 13 EVA-2 Timeline

LM Ascent Stage DeorbitLunar Surface Communications NetworkManned Space Flight Network (Apollo 13)Apollo 13 Primary Landing AreaPrimary Landing Area Recovery Force Voice CallsRecovery Force SupportPrime Crew of Seventh Manned Apollo Mission

Page9

14

151718191920

2124

2526333436373842


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TableLIST OF TABLES

TitleApollo 13 Launch Windows

Apollo 13 Sequence of Major EventsApollo 13 TV ScheduleApollo 13 Weight Summary

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101112

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SUMMARY OF APOLLO/SATURN MISSIONSMissionAS-201

AS-203

AS-202

APOLM 4

APOLLO 5

LaunchDate2/26/66

l/5/66

B/25/66

11/g/67

l/22/68

LaunchVehicleSA-201

SA-203

SA-202

SA-501

SA-204

PayloadCSM-009

LH2 inS-WE

cm-011

CSM-017LTA-1OR

r&l-lSLA-7

DescriptionLaunch vehicle end CSMdevelopment. Test ofCSM subsystems and of thespace vehicle. Demon-stration of reentry ade-quacy of the CM at earthorbital conditions.Launch vehicle development.Demonstration of controlof LH2 by continuous vent-ing in orbit.Launch vehicle end CSMdevelopment. Test of CSMsubsystems and of thestructural integrity andcompatibility of the spacevehicle. Qemonstratio" ofpropulsion and entry con-trol by G&N system. Demon-stration of entry at 28,500fps.Launch vehicle and space-craft development. Demo"-stration of Saturn V LaunchVehicle performance and ofCM entry at lunar returnvelocity.LM development. Verifiedoperation of LM subsystems:ascent and descent propul-sion systems (includingrestart) and structures.EValUatiOn Of LM Staging.Evaluation of S-IVB/ILl or-bital performance.

MissionAPOLLO 6

APOLLO 7

APOLLO 8

APOLLO 9

APOLLO10

APOLLO 11

APOLLO 12

LaunchDate4/4/68

10/11/68

12/21/68

3/3/69

5/18/69

7/16/69

11/14/69

LaunchvaiTEieSA-502

SA-205

SA-503

SA-504

SA-505

SA-506

SA-507

PayloadCM-020SM-014LTA-2RsLA-9CM-101SM-101.%A-5CM-103SM-103LTA-BS&A-11CM-104SH-104LM-3SIA-12CM-106SM-106LM-4SLA-13

CM-107SM-107LM-5SLA-14CM-108SM-108LJ4-6SLA-15

DescriptionLaunch vehicle and space-craft development. Demon-stration of Saturn VIaunch Vehicle perfor mance.Manned CSM operations.Duration 10 days 20 hours.Lunar orbital mission.Ten lunar orbits. Missionduration 6 days 3 hours.Manned CSM operations.Earth orbital mission.Manned CSM/LM operations.Duration 10 days 1 hour.Lunar orbital mission.Manned CSWLM ooerations.Evaluation'of ti perform-ance in cislunar and lunarenvironment, followinglunar landing profile.Mission duration 8 days.First manned lunar lan dingmission. Lunar surfacestay tim e 21.6 hours.Mission duration 8 days3 hours.Second manned lunar landingmission. Demonstration ofpoint landing capability.DeDhVWZt Of ALSEP I.Su+or III investigation.Lunar Surface stay time31.5 hours. Two dual EVA's(15.5 manhours). 89 hours inlunar orbit (45 orbits).Mission duration 10 days4.6 hours.


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NASA OMSF PRIMARY MISSION OBJECTIVES FOR APOLLO 13PRIMARY OBJECTIVES

. Perform selenological inspection, survey, and sampling of materials in apreselected region of the Fra Mauro Formation.

. Deploy and activate an Apollo Lunar Surface Experiments Package (ALSEP).

. Develop mans capability to work in the lunar environment.

. Obtain photographs of candidate exploration sites.

Apol lo Program DirectorManned Space Flight

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DETAILED OBJECTIVES AND EXPERIMENTSSPACECRAFT DETAILED OBJECTIVES AND EXPERIMENTS. Contingency Sample Collection.. Apollo Lunar Surface Experiments Package (ALSEP III) Deployment.. Selected Sample Collection.

Lunar Field Geology (S-059).Photographs of Candidate Exploration Sites.Evaluation of Landing Accuracy Techniques.Television Coverage.EVA Communica tion System Performance.Lunar Soil Mechanics.Selenodetic Reference Point Update.Lunar Surface Close-Up Photography (S- 184).

Thermal Coating Degradation.CSM Orbital Science Photography.Transearth Lunar Photography.Solar Wind Composition (S-080).Extravehicular Mobility Unit Water Consumption Measurement.Gegenschein From Lunar Orbit (S-178).

Dim Light Photography.CSM S-band Transponder (S- 164).Down1 ink Bistatic Radar (VHF Only) (S- 170).

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LAUNCH VEHICLE DETAILED OBJECTIVES. Impact the expended S-lVB/IU on the lunar surface to excite ALSEP I.. Determine actual S-lVB/lU point of impact.

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LAUNCH COUNTDOWN AND TURNAROUND CAPABILITY, AS-508COUNTDOWNCountdown for launch of the AS-508 Sp ace Vehicle for the Apollo 13 Mission willbegin with a precount starting at T-94 hours during which launch vehicle and spacecraftcountdown activities will be conducted independently. Official coordinated spacecraftand launch vehicle countdown will begin at T-28 hours.

SCRUB/TURNAROUNDA scrub is a termination of the countdown. Turnaround is the time required to recycleand count down to launch (T-O) in a subsequent launch window assuming no serialrepair activities are required. The scrub/turnaround plan will be placed in effectimmediately following a scrub during the countdown.For a hold that results in a scrub prior to T-22 minutes , turnaround procedures areinitiated from the point of hold. Should a hold occur from T-22 m inutes (S-II startbottle chilldown) to T-16.2 seconds (S-IC f orward umbilical disconnect), then arecycle to T-22 minutes, a hold, or a scrub is possible under conditions stated in theLaunch Mission Rules. A hold between T-16.2 seconds and T-8.9 seconds (ign ition)could result in either a recycle or a scrub depending upon the circumstances. Anautomatic or manual cutoff after T-8.9 seconds will result in a scrub.30-DAY SCRUB/TURNAROUND

A 30-day turnaround capability exists in the event that a scrub occurs and there is nolaunch window available within the 24 or 48-hour turnaround capability.In the event of a 30-day scrub/turnaround a new countdown will be started at thebeginning of precount. The Flight Readiness Test (FRT) and Countdown DemonstrationTest (CDDT) will not be rerun.48-HOUR SCRUB/TURNAROUNDThe maximum scrub/turnaround time from any point in the launch countdown up toT-8.9 seconds is 48 hours. Th is maximum time assumes no serial repair activities arerequired and it provides for reservicing all space vehicle cryogenics.24-HOUR SCRUB/TURNAROUNDA 24-hour turnaround capability exists as late in the countdown as T-8.9 seconds.This capability depends upon having sufficient spacecraft consumables margins aboveredline quantities stated in the Launch Mission Rules for the period remaining to thenext launch window. Only one 24-hour scrub/turnaround can be accomplished.

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FLIGHT MISS ION DESCRIPTION

LANDING SITEThe landing site of the Apollo 13 Mission is a point 34OS latitude, 17O29 W longitudein the Fra Mauro Formation. The Fra Mauro Formation, an extensive geologic unitcovering large portions of the lunar surface around Mare Imbrium, has been interpretedas the ejecta blanket deposited during the formation of the lmbrium basin. Samplingof the Fra Mauro Formation may provide information on ejecta blanket formation andmodification, and yield samples of deep-seated crustal material giving information onthe composition of the lunar interior and the processes active in its formation. Agedating the returned samples should establish the age of premare deep-seated materialand the age of the formation of the lmbrium basin and provide important points on thegeologic time scale leading to an understanding of the early h istory of the moon.LAUNCH WINDOWS

The launch windows for Fra Mauro are shown in Table 1.

LAUNCH DATE

April 11, 1970**May 9, 1970 (T-24)**May 10, 1970 (T-O)**May 11, 1970 (T+24)

TABLE 1

APOLLO 13 LAUNCH WINDOWS SUNOPEN CLOSE ELEVATION ANGLE*

14: 13 17:37 9.9O13:25 16:43 7.813:35 16:44 7.813:32 16:38 18.5

NOTE: Only one scrub/turnaround is feasible for May. April times are EST; allothers are EDT.

* These values are subject to possible refinement.** The addition of the T-24 hour and T+24 hour windows to the optimum T-O window

provides increased flexibility in that all three opportunities are available forchoice.

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LAUNCH OPPORTUNITIESThe three opportunities established for May --in case the launch is postponed from11 Apri I --provide, in effect, the flexibility of a choice of two launch attempts. Theoptimum May launch window occurs on 10 May. The 3-day window permits a choiceof attempting a launch 24 hours earlier than the optimum window and, if necessary, afurther choice of a 24-hour or 48-hour recycle. It also permits a choice of making thefirst launch attempt on the optimum day with a 24-hour recycle capability. The 9 Maywindow (T-24 hrs.) requires an additional 24 hours in lunar orbit before initiatingpowered descent to arrive at the landing site at the same time and hence have the samesun angle for landing as on 10 May. Should the 9 May window launch attempt bescrubbed, a decision will be made at that time, based on the reason for the scrub,status of spacecraft cryogenics and weather predictions, whether to recycle for 10 May(T-O hrs. ) or 11 May (T+24 hrs .) If I aunched on 11 May, the flight plan will be similarfor the 10 Mayomission but the sun elevation angle at lunar landing will be 18.5instead of 7.8

HYBRID TRAJECTORYThe Apollo 13 Mission will use a hybrid trajectory that retains most of the safety featuresof the free-return trajectory but without the performance limitations. From earth orbitthe spacecraft will be initially injected into a highly eccentric elliptical orbit (peri-cynthion of approximately 210 nautical miles (NM), which has a free-return character-istic, i.e., the spacecraft can return to the earth entry corridor without any furthermaneuvers. The spacecraft wi II not depart from the free-return ellipse unti I after theLunar Module (LM) has been extracted from the launch vehicle and can provide a pro-pulsion system backup to the Service Propulsion System (SPS). Approximately 28 hoursafter translunar injection (TLI), a midcourse maneuver wil I be performed by the SPS toplace the spacecraft on a lunar approach trajectory (non-free-return) having a peri-cynthion of 60 NM.The use of a hybrid trajectory will permit:

Daylight launch/Pacific injection. This allows the crew to acquire thehorizon as a backup attitude reference during high altitude abort, provideslaunch abort recovery visibility, and improves launch photographic coverage.

Desired lunar landing site sun elevation. The hybrid profi le faci Ii tatesadjustment of translunar transit time which can be used to control sunangles on the landing site during lunar orbit and at landing.

-

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M-932-70-1 3

Increased spacecraft performance. The energy of the spacecraft on a hybridlunar approach trajectory is relatively low compared to what it would be ona full free-return trajectory, thus reducing the differential velocity (AV)required to achieve lunar orbit insertion.Improved communication flexibility. This permits adjustment of the time ofpowered descent initiation (PDI) to occur within view of a 210-foot groundantenna.

LIGHTNING PRECAUTIONSDuring the Apollo 12 Mission, the space vehicle was subjected to two distinct electri-cal discharge events. However, no serious damage occurred and the mission proceededto a successful conclusion. Intensive investigation led to the conclusion that no hard-ware changes were necessary to protect the space vehicle from similar events. ForApollo 13 the Mission Rules have been revised to reduce the probability that the spacevehicle will be launched into cloud formations that contain conditions conducive toinitiating similar electrical discharges although flight into all clouds is not precluded.

- FLIGHT PROFILELaunch Through Earth Parking OrbitThe AS-508 Space Vehicle for the Apollo 13 Mission is planned to be launched at14:13 EST on 11 Apri I 1970 from Launch Complex 39A at the Kennedy Space CenterFlorida, on a flight azimuth of 72. The Saturn V Launch Vehicle (LV) will insertthe S-IVB/lnstrument Unit (lU)/LM/CSM into a 103-NM, circular orbit. The S-lVB/IUand spacecraft checkout will be accomplished during the orbital coast phase. Figure 1and Tables 2 through 4 summarize the flight profile events and space vehicle weight.Translunar InjectionApproximately 2.6 hours after liftoff, the launch vehicle S-IVB stage will be reignitedduring the second parking orbit to perform the translunar injection (TLI) maneuver,placing the space vehicle on a free-return trajectory having a pericynthion of approxi-mately 210 NM.

Transl unar Coast

.-

The CSM will separate from the S-lVB/IU/LM approximately 4 hours Ground ElapsedTime (GET), transpose, dock, and initiate ejection of the LM/CSM from the S-lVB/IU.During these maneuvers, the LM and S-lVB/lU will be photographed to provide engineer-ing data.

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Yy\2TRANSEARTHASCENT STAGE

TPI -1-i x 61 NM

CMSM SEPARATIONLM INSERTION ;.9 x 44 - NM UKtSIII

CSM\ /&

EIVB RESTARTDURING 2ND OR3RD ORBIT

S-IV6 2ND BURNTRANSLUNAR INFRFF-RFTIIRN T

SPLASHDCOVERYS-IVB AP! TRAJECTORY. TRA IFrTnRV

- , _ . . . LECTO RY/ I

TRIM BURN(S)SC SEPARATION, S-IVB RESIDUALTRANSPOSITION, PROPELLANTDOCKING, & EJECTION ouw (IMPACT)AND SAFING

-LAUNCH INITIATION (REV 14)

I I\nYL I I\ ITRANSFERI: (I.2 ORBITS)::

,CSM ORBITCIRCULARIZ(REV 12) 60LOI CSM/LM60x170 NM(XT 2 ORBIT

-DOICSMILM

ATION;cl

5)

S-IVB AP+EVASIVE MANEUVER

APOLLO13 FLIGHTPROFILE


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EVENTLAUNCHTLIMIDCOURSECORRECTION-lMIDCOURSEORRECTION-ZMIDCOURSE CORRECTIONS (AS REQ'D)LUNAR ORBIT NSERTIONLOI)DESCENTORBIT NSERTIONCSM/LM UNDOCKING & SEPARATIONCSMORBITCIRCULARIZATIONPOWEREDDESCENT INITIATION (PDI)LANDING RAMAUROEVA-lEVA-2ASCENT (LM LIFTOFF)DOCKINGLM IMPACTURNTRANSEARTHINJECTIONTEI)ENTRY INTERFACELANDING

TABLE 2APOLLO 13 SEQUENCEOF MAJOR VENTS

GET HR:MIN0o:oo

2:3511:41 AS REQ'D30:41 2 (SPS)77:2581:4599:16

100:35103:31103:42108:OO127:45137:09140:45144:32167:29240:50241:03

BURN DURATION-SEC. (SYSTEM)

358 (S-IVB)

357 (SPS)23 (SPS)6 (SM RCS)4 (SPS)687 (DPS)

528 (APS)75 (LM RCS)135 (SPS)

REMARKSPAD 39A, 4/11/70, 14:13 ESTPACIFIC NJECTIONNOMINALLYEROHYBRID RANSFER74.9-HR. TRANSLUNAR COAST60 x170-NM ORBIT8 X ~&NM ORBIT CSM/LM)SOFT UNDOCKING53 x 62-NM ORBIT REV. 12)LANDING O1T UPDATESEA = 7.1 4/15, 21:55 EST4 HR. PLANNED4 HR. PLANNED4/17, 7:22 ESTCSM-ACTIVEASCENT STAGE IMPACT T 145:0073.6-HR. TRANSEARTH COASTVELOCITY 6, 129 FPSPACIFIC, 4/21, 15:16 EST


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TABLE 3

DAYSATURDAY

SATURDAY

SUNDAY

TUESDAY

WEDNESDAY

THURSDAY

THURSDAY

FRIDAY

SATURDAY

SATURDAY

MONDAY

DATEAPRIL 11

APRIL 11

APRIL 12

APRIL 14

APRIL 15

APRIL 16

APRIL 16

APRIL 17

APRIL 18

APRIL 18

APRIL 20

EST15:49

17:28

20:28

00:13

14:03

02:23

22:03

lo:36

12:23

14:13

19:58

APOLLO 3 TV SCHEDULE

GET01:36

03:15

30:15

58:00

95:50

108:lO

127:50

140:23

166:lO

168:OO

221:45

DURATION05 MIN

1 HR 08 MIN

30 MIN

30 MIN

15 MIN

3 HR 52 MIN6 HR 35 MIN

12 MIN40 MIN25 MIN

15 MIN

ACTIVITY/SUBJECTCOLOR PHOTOSOF EARTH

TRANSPOSITION & DOCKING

SPACECRAFT NTERIOR (MCC-2)

INTERIOR & IVT TO LM

FRA MAURO

LUNAR SURFACE (EVA-~)

LUNAR SURFACE (EVA-2)

DOCKINGLUNAR SURFACE

LUNAR SURFACE (POST TEI)EARTH & SPAC ECRAFT NTERIOR

VEHCSM

CSMCSMCSM

CSM

LM

LM

CSMCSMCSMCSM

STAKSC

GDS

GDS

GDS

MAD

GDS/HSK

GDS

MAD

MAD*

MAD*GDS

* RECORDEDONLY


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TABLE 4APOLLO 13 WEIGHT SUMMARY(We

s-IC/S-IIInterstageS-II StageS-II/S-IVBInters tageS-IVB StageInstrument Unit

11,46478,0508,100

25,0504,482

Spacecraft-LMAdapterLunar ModuleService ModuleCommand ModuleLaunch EscapeSystem

jht in PoundsTOTALEXPENDABLES

4,746,870---996,960 1,075,OlO 92,523

m-m 8,100 mm -236,671---

I-INHLSEPARATIONTOTAL WEIGHT WEIGHT5,034,870 363,403

11,464 mm-

261,7214,482

Launch Vehicle at Ignition 6,395,647

35,526--- 14,0449,915

10,53212,572

9,012

23,56840,567

---

4,04433,483 33,94151,099 **14,07612,572 **11,269

9,012(Landing)

---

Spacecraft at Ignition 110,210 1Space Vehicle at Ignition 6,505,857S-IC Thrust Buildup (-)84,598Space Vehicle at Liftoff 6,421,259Space Vehicle at Orbit Insertion 299,998

* CSM/LM Separation** CM/SM Separation

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M-932 -70-l 3

At 4.2 hours GET the S-IVB/IU will begin a series of programmed and ground-commandedoperations which will alter the LV trajectory so that the S-lVB/lU will impact the lunarsurface at the desired point providing a known energy source for the Apollo 12 ALSEPseismology equipment. The first is a programmed Auxiliary Propulsion System (APS)ullage mo tor retrograde burn evasive maneuver to provide initial launch vehicle/spacecraftseparation to prevent recontact of the two vehicles. Second, by a combination of pro-grammed liquid oxygen (LOX) dumping the S-lVB/lU will be placed on a lunar impacttrajectory. A second APS ul lage motor burn wi I I be ground commanded at approximately6.0 hours GET. The burn duration and attitude will be determined in real-time basedontrajectory data. This burn is intended to place the S-IVB/IU on the trajectory for lunarimpact at the desired point. A third APS ullage motor burn, also ground commanded, willbe performed if necessary to refine the S-lVB,/lU trajectory. This burn will occur approxi-mately 9oO hours GET. The desired impact will be within 350 kilometers of the targetpoint, 3 S. latitude, 30W. longitude. The impact will occur while the CSM/LM is onthe backside of the moon. Later, the crew will photograph the S-IVB/IU target impactarea. It is desired that postflight determination of actual impact be within 5 kilometersin distance and 1 second in time.The spacecraft will be placed on a hybrid trajectory by performing a Service PropulsionSystem (SPS) maneuver at the time scheduled for the second Midcourse Correction (MCC),approximately 30.6 hours GET. The CSM/LM combination will be targeted for a peri-cynthion altitude of 60 NM and, as a result of the SPS maneuver, will be placed on anon-free-return trajectory. The spacecraft will remain within the LM Descent PropulsionSystem (DPS) as well as the SPS return capability.Earth weather will be photographed for a 3 to 4-hour period beginning at approximately7 hours GET. Lunar photgraphy may be performed, at the crews option, at 56 and 75hours GET. MCCs will be made as required, using the Manned Space Flight Network(MSF N) for navigation.Lunar Orbit InsertionThe SPS will insert the spacecraft into an initial lunar orbit (approximately 60 x 170 NM)at 77.6 hours GET (Figure 2). The spacecraft will remain in a 60 x 170-NM orbit forapproximately two revolutions.Descent Orbit InsertionAfter two revolutions in lunar orbit, the SPS will be used to insert the spacecraft in a60 x 8-NM descent orbit. During the 4th revolution, the Command Module Pilot (CMP)will photograph the candidate exploration site Censorinus from the CSM at low altitudewhile the Commander (CDR) and Lunar Module Pilot (LMP) enter the LM for checkoutand housekeeping. The crew will use approximately six revolutions for eat and restperiods, and then wi I I prepare the LM for separation and powered descent.

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APPROACHHYPERBOLA

8 x 60 NM(12 REVS)/

EARTH\

LUNARORBIT INSERTIONAND DESCENTORBITNSERTIONGEOMETRY


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A soft undocking will be made during the 12th revolution. Spacecraft separation willbe executed by the Service Module Reaction Control System (SM RCS) with the CSMradially below the LM (Figure 3).

LM DESCENT ORBITAL EVENTS

CSM ClRCULARlZATlON (REV 12)(53 x 62 I1M) I

LM DESCENT ORBIT(8 x 60 !IM)

-. UNDOCKlNG ANDSEPARATION (REV

/ .SUN

EARTH

Fig. 3CSM Lunar Solo (Pre-Rendezvous)

12)

During the 12th revolution (after undocking) the CSM will perform a circularizationmaneuver to a near-circular 60-NM orbit. The CSM will photograph selected sites that

- will include the candidate exploration site Censorinus. Lunar orbital science photographyand dim light photography of zodiacal light, solar corona, and gegenschein will beperformed.

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Lunar Module DescentDuring the 14th revolution the DPS will be used for the powered descent maneuver,which wil I start approximately at pericynthion. The vertical descent portion of anautomatic landing during the landing phase will start at an altitude of about 100 feetand will be terminated at touchdown on the lunar surface (Figure 3). The crew may

s elect to take over manually at an altitude of about 500 feet or below. Return toautomatic control is a newly added capability of the LM guidance computer.During descent, the lunar surface will be photographed from the LMPs window torecord LM movement and surface disturbances and to aid in determining the landedLM location.Lunar Surface Operations

A summary of the lunar surface activities is shown in Figure 4.PostlandinaImmediately upon landing, the LM crew will execute the lunar contact check-list and reach a stay/no-stay decision. After reaching a decision to stay, theInertial Measurement Unit (IMU) will be aligned, the Abort Guidance System(AGS) gyro will be calibrated and aligned, and the lunar surface will bephotographed through the LM window. Following a crew eat period al I looseitems not required for extravehicular activity (EVA) will be stowed.

EVA-lThe activity timeline for EVA-l is shown in Figure 5. Both crew members willdon helmets, gloves, Portable Life Support Systems (PLSS), and Oxygen PurgeSystems (OPS) and th e cabin will be depressurized. The CDR wil I move throughthe hatch, deploy the Lunar Equipment Conveyor (LEC), and move to theladder where he will deploy the Modularized Equipment Stowage Assembly(MESA), Figure 6, which initiates television coverage from the MESA. He willthen descend the ladder to the lunar surface. The LMP will monitor andphotograph the CDR using a still camera (70mm Hasselblad Electric Camera)and the lunar geologic exploration sequence camera (16mm Data AcquisitionCamera).Environmental Familiarization/Contingency Sample Collection - After steppingto the surface and checking his mobility, stability, and the ExtravehicularMobility Unit (EMU), the CDR will collect a contingency sample. This willmake it possible to assess the nature of the lunar surface material at theApollo 13 landing site in the event the EVA were terminated at this point.The sample will be collected by quickly scooping up a loose sample of the

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,

TOUCHDOWN

t-rOST- EATLAND,APOLLO13

LUNAR SURFACE CTIVITIES SUMMARY103:42 GET (2155 EST)

EVA 1STPREP, EVA2 HR, 4 HR,

rc/O

;;;- EAT PLSS SLEEPCHECK- EVfOUT DEBRIEF-ING142HRi HRn0,5 HR, 916 HR,IIIIIIIIIIIIIIIIGDS IIIIIIIIIIIIIIIIIIIIIIIIJIIIIIIIIIIIIc

EAT 1 EVAI REP,I.,2 H , 2 HR,I 2NDEVA POST- / LAUNCHPREP,4 HRu EVA C/O& DEBRIEF-ING215HR, 3 HR,lllllllllllllllllllllllllllllllllllllllllll~DS llllllllllllllllllllllnllllllllllllllllllllllllllllll 7n-.ccl.P J 3315 HR, TOTALTIMEEVAEXTENDABLEO 5 HOURSASEDONA REALTIME SSESSMENTF REMAININGONSUMABLES,

LIFTOFF m


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APOLLO 13 EVA-I TIMELINE

CDR

LMP

IES;"p I

S-BANDIPHOTODEPLOY

I ITRANSFERS DEP. iALSEP ;;FLINSP. OFF PHOTOAND-I,h TV PANSS TI I:1

30 40

ANT. SW.TRANSFERS

50DEP. INSP. OFF- FUEL SRC 1LOAD UP

CENTRAL STATION ACTIVATICN/ ,ALSEP DEPLOY ASSEMBLEDRILL DEPLOY COREDRILL BORE HOLE 1

lt40 160 2tOo 2t10 2+20 2+30 2+40 2t50 3tOo

n-.M.0-l

CDR FINAL SELECTED SRC PACK AND CLOSE OUT I I--..

1LMP

3tOo

TRANSFERSLSEP SAMPLEPHOTOSCLOSLOUTDEBRIS SELECTED OFF-LOAD SRC 2 .INGRESSCLEAN SAMPLE DEPLOY SWC TRANSFERSJP .INGRESS PREP.

3+10 3+20 3+30 3t4D 3+50

JINGR~ !EvA EXTENSION(N1tOo 4t10


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DEPLOYED MODULARIZEDEQUIPMENT STOWAGE ASSEMBLY

3/27 PO

Fig. 6

lunar material (approximately 2pounds), sealing it in a Contin-gency Sample Container, andplacing the sample in the Equip-ment Transfer Bag (ETB), along withthe lithium hydroxide (LiOH)canisters and PLSS bat.teries forlater transfer into the LM using theLEC. The LMP will transfer the70mm cameras to the surface withthe LEC. The LMP wi II then descendto the surface.S-band Antenna Deployment -The S-band antenna wil I be re-moved from the LM and carried tothe site where the CDR will erectit, as shown in Figure 7, connectthe antenna cable to the LM, andperform the required alignment.

DEPLOYED S-BAND ANTENNA

Fig. 7Page 19


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Lunar TV Camera (Color) Deployment - While the CDR deploys the S-bandantenna, the LMP will unstow the TV camera and deploy it on the tripodapproximately 50 feet from the LM. The LMP will then ingress the LM toactivate and verify TV transmission with the Mission Control Center.

The contingency sample and other equipment will be transferred into the LMand the 16mm lunar geologic exploration camera transferred to the surface.The LMP will egress again, leaving the hatch slightly ajar, and descend tothe surface. The CDR wil I photograph the contingency sample area and theLMP egress. Following this, the American flag wil I be deployed.LM Inspection - After repositioning the TV to view the scientific equipmentbay area, the LMP will inspect and photograph the LM footpads and quadrantsI, II, Ill, and IV with an EMU-mounted 70mm camera. Concurrently the CDRwill also inspect and photograph the LM. The Apollo Lunar Surface Close-UpCamera (ALSCC) will be removed from the MESA and placed down sun.ALSEP Deployment - After offloading ALSEP from the LM, the RadioisotopeThermoelectric Generator (RTG) will be fueled (Figure 8 ), the ALSEP packageswill be attached to a one-man carry bar for traverse in a barbell mode, andthe TV will be positioned to view the ALSEP site. The hand tools will beloaded on the hand tool carrier. While the CDR obtains TV and photographicpanoramic views from the site, the LMP will unload Sample Return Containernumber 1 (SRC 1) and remove the ALSCC. The CDR and LMP wi II then carrythe ALSEP packages, hand tool carrier, and Apollo Lunar Surface Drill (ALSD)to the deployment site approximately 500 feet from the LM. The crew willsurvey the site and determine the desired location for the experiments. Thefollowing individual experiment packages will then be separated, assembled,and deployed to respective sites in the arrangement shown in Figure 9.

Fig. 8

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-Y=: EVA-l EQUIPMENTDEPLOYMENT\2TV PLACEMENT INITIAL TV PLACEMENTFOR ALSEP BY LMP

ALSEP DEPLOYMENT SITE TV POSITION FORSEQ. BAY ACTIVITY

SAMPLE COLLECTION ALSEP OFFLOADAREA OF ACTIVITY

CCGE

n-.ul.9

LEGEND: LMP - Lk PI-LOTCDR - COMMANDERALSEP - APOLLO LUNAR SURFACE EXPERIMENTS PACKAGESEQPSE - SCFENTIFIC EQUIPMENT- PASSIVE SEISMIC EXPERIMENTHFE - HEAT FLOW EXPERIMENTCPLEECCGE - CHARGED PARTICLE LUNAR ENVIRONMENT EXPERIME-WT- COLD CATHODE GAUGE EXPERIMENTRTG - RADIOISOI;oPIC THERMOELECTRICGENERATOR


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Passive Seismic Experiment (S-031) - The CDR will deploy and set up thePassive Seismic Experiment package with its thermal cover.

Charged Particle Lunar Environment Experiment (S-038) - The CDR will deployand orient the Charged Particle Lunar Environment Experiment package whilethe LMP is assembling the ALSD and drilling the first hole for the heat flowexperiment.Cold Cathode Ion Gauge (S-058) - The CDR will deploy and orient the ColdCathode Ion Gauge.Heat Flow Experiment (S-037) - The LMP will assemble the battery-poweredALSD and will drill two three-meter deep holes using hollow-center bore stems.Each bore stem will be left in place as an encasement into which the heat flowprobes are inserted.

ALSEP Central Station - The CDR will I evel and align the central station whichincludes deployment of the sun shield. At this time the LMP will be drillinathe second hole for the Heat F low Experiment. The CDR will then assembleand align the ALSEP antenna. The CDR will activate the central station andphotograph the ALSEP layout while the LMP is implanting the Heat FlowExperiment probes.Core Sampling - The CDR will assist the LMP in modifying the ALSD, collectingcore samples and photographing the operation.

Selected Sample Collection - The crew will begin the return traverse andinitiate collection of the selected samples. At the return to the LM, sampleswill be weighed and, with the core stems, stowed in SRC 1 and the SRC will besealed. SRC 2 will be offloaded and SRC 1 will be transferred into the LM.Solar Wind Composition Experiment (S-080) - The four-square-foot panel ofaluminum foil will be deployed by the LMP.EMU cleaning and ingress into the LM will be accomplished by the LMP andthe CDR. EVA-l will terminate when the LM cabin is repressurized.Post-EVA 1 Ooerations

After configuring the LM systems for Post-EVA-l operations, the PLSSs willbe recharged. This includes filling the oxygen system to a minimum pressureof 875 pounds per square inch, filling the water reservoir, and replacing thebattery and LiOH canister. A one-hour eating period is scheduled betweenthe beginning and end of the PLSS recharge operations. The PLSSs and OPSswill be stowed, followed by a 9-1/2-h our rest period and another eat period.

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M-932-70-13EVA-2The LM will be configured for EVA activities and the CDR will egress. TheLMP will again monito r and photograph the CDR and then transfer cameraequipment in the ETB to the CDR with the LEC. The LMP will descend to thelunar surface leaving the LM hatch slightly ajar. A summary of EVA-2activities is shown in Figure 10.Sample Collection and Camera Calibration - The crew will collect a thermalsample and a sieve sample from near the LM and a contaminated sample fromunder the LM. They will calibrate their cameras by photographing a specialcontrast chart on the hand tool carrier and then will position the TV camera forthe field geology traverse.

Lunar Field Geology (S-059) Traverse - Both crewmen will conduct the fieldgeology traverse, which is planned in detai I prior to launch. Additional supportand real-time planning will be provided from the ground based on features of thelanding site obtained from crew descriptions and TV. Traversing outbound fromthe LM, the crew will obtain Close-Up Stereo Camera photos of selected areas.They will take panoramic photographs of the lunar surface and will use a specialpolarizing filter to photograph selected features. They will obtain subsurface.samples, core camples, and surface samples. Special gas analysis, environmental,and magnetic lunar surface samples will be collected. Approximately l/2 milefrom the LM the CDR will dig a two-foot-deep trench for lunar soil mechanicsevaluation. The LMP will collect a core sample and a special environmentalsample from the trench and obtain photographic data of the boot prints in thetrench material. Gas analysis and magnetic samples will also be obtained inthe vicinity of the trench.

The crew will begin traversing inbound to the LM by a different route toobtain additional documented samples as performed on the outbound traverse.A typical documented sample procedure will include locating the gnomon upsun of the sample site, photography of the site and the sample, description ofthe sample, and stowage of the lunar surface material in a sample bag. Thetypical core sampling will consist of placing the gnomon up sun and photographyof the sample site cross sun, driving the core tube into the surface, recovery andcapping the sample within the tube.Upon return to the LM, the CDR will offload the samples into SRC 2. TheLMP will reposition the TV, collect a soil mechanics sample, and then takedown and roll up the Solar Wind C omposition experiment and place it and theClose-Up Stereo Camera magazine in the SRC. He will then close and sealthe SRC. The LMP will clean his EMU, ingress into the LM, and hook up theLEC. The CDR and LMP will utilize the LEC to transfer samples and equipmentinto the LM. The CDR will clean his EMU, ascend into the LM, jettison theLEC and ingress. The LM will be repressurized, terminating EVA-2. Equip-ment and samples will be stowed and preparations made for equipment jettison.The LM will be depressurized, equipment jettisoned, and the LM repressurized.

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Y2:-a0CDR

LMP0

PLSSCHECKOUT

PLSS CAMERACHECKOUT TRANS.

EGRESSCAMERATRANS.

APOLLO 13 NA-2 TIMELINE

SRC TRAV. THERM SIEVE COLOROPEN PREP, SMPLE SMPLE CHART GEOLOGY RAVERSEALHT N pT LEGRS. TRAV. THERM SIEVE COLORF';';. SMPLE SMPLE CHART Q p GEOLOGY RAVERSF

T R

BIG TRENCH2 SPECIAL ENVIR. SAMPLES(0ID CDR GAS ANALYSIS SAMPLE CORE TUBES2 MAGNETIC SAMPLE

LMP1+40 1+50 2too 2t10 2t20 2+30 2+40 260 3too

CDR SRC PACK AND SRC & EQUIP. INGRESS EVA EXTENSIONCLOSEOUTACT. TRANSFERS,n-. SOIL SAMPLE(I) INGRESS EVA EXTENSION. LMP RETRIEVE SWCCLOSEOUTACT. TRANSFERS"0 I m I I3tOo 3+10 3+20 3+30 3t40 360 4tOO 4+10 4t20


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Lunar Module AscentThe LM ascent (Figure 11) will begin a fter a lunar stay of approximately 33.5hours. The Ascent Propulsion System (APS) povvered ascent is divided into twophases. The first phase is a vertical rise, which is required to achieve terrainclearance, and the second phase is orbit insertion. After orbit insertion theLM will execute the coelliptic rendezvous sequence wh ich nominally consistsof four major maneuvers: concentric sequence initiation (CSI), constantdelta height (CDH), terminal phase initiation (TPI), and terminal phasefinalization (TPF). A nominally zero plane change (PC) maneuver will bescheduled between CSI and CDH, and two nominally zero midcourse correctionmaneuvers will be scheduled between TPI and TPF; the TPF maneuver is actuallydivided into several braking maneuvers. All maneuvers after orbit insertionwill be performed with the LM RCS. Once docked to the CSM, the two crew-men will transfer to the CSM with equipment, lunar samples, and exposed film.Decontamina tion operations will be performed, jettisonable items will beplaced in the Interim Stowage Assembly and transferred to the LM, and the LMwill be configured for deorbit and lunar impact.

LM ASCENT THROUGH DOCKING,~~~=-~-=~~~Rendezvous radar tracking

l --- MSFN tracking

1 . LiftoffInsertion. LM3. CSI4. PC5. CDt -I

6. TPI Earth7. MC-l8. MC-29. Begin Braking10. Begin Stationkeeping- 11. Docking

CSM parking orbit

EVENT

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-Sun

Fig. 11


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LM Ascent Stage Deorbit

CSM

The ascent stage will be deorbited (Figure 12), during the 35th revolution,for lunar surface impact between the ALSEP I of Apollo 12 and the newlydeployed ALSEP Ill, to provide a known energy source to produce signals forrecording by the seismic experiments. The CSM will be separated radiallyfrom the ascent stage with a SM RCS retrograde burn approximately 2 hoursafter docking to the CSM. Following the LM jettison maneuver, the CSMwill perform a pitchdown maneuver. The LM deorbit maneuver will be aretrograde RCS burn initiated by ground control and the LM will be targetedto impact the lunar surface approximately 36.5 NM west northwest of theApollo 13 landing site. The ascent stage jettison, ignition, and impactedlunar surface area will be photographed from the CSM.

LM ASCENTSTAGEDEORBIT

lMPACT( 36.5 rq WEST NORTHWEST /

OF LANDING SITE) /

JETTISONRADIAL)EARTH

Fig. 12

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CSM Lunar Orbit OperationsThe CSM wi II execute an orbital plane change, soon after MSFN acquisition of signalin the 35th revolution, for approximately 9 hours of lunar reconnaissance photography.High resolution vertical and oblique topographic photography, stereo strip photography,science photography and landmark tracking will be performed. The Censorinus,Descartes, and Davy Rille sites are of special photographic interest as candidate futurelanding sites.Transearth lniection and CoastAt the end of the 46th revolution, and approximately 10 minutes prior to MSFNacquisition of signal, the SPS wil I be used to inject the CSM onto the transearthtrajectory. Th e return flight duration will be approximately 73 hours (based on an11 April launch) and the return inclination (to the earths equator) will not exceed40 degrees. Md course corrections wil I be made as required, using the MSFN fornavigation.

Entry and Landing-Prior to atmospheric entry, the CSM will maneuver to a heads-up attitude, theCommand Module (CM) will jettison the SM and orient to the entry attitude.The nominal range from entry interface (El) at 400,000 feet altitude to landing willbe approximately 1250 NM. Earth landing will nominally bel34S latitude and 15730W longitude (based on an 11 April

241 hours after liftoff.

in the Pacific Ocean atlaunch) approximately

Crew Recovery and QuarantineFollowing landing, the Apollo 13 crew will don the flight suits and face masks passedin to them through the spacecraft hatch by a recovery swimmer wearing standardscuba gear. Integral Biological Isolation Garments (BIGS) will be available for usein case of an unexplained crew illness. The swimmer will swab the hatch and adjacentareas with a liquid decontamination agent. The crew will then be carried by helicopterto the recovery ship where they will enter a Mobile Quarantine Facility (MQF) and allsubsequent crew quarantine procedures will be the same as for the Apollo 11 and 12Missions.

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CM and Data Retrieval OperationsAfter flight crew pickup by the helicopter, the CM wil I be retrieved and placed ona dolly aboard the recovery ship. The CM will be mated to the MQF, and the lunarsamples, film, flight logs, etc., will be retrieved and passed out through a decon-tamination lock for shipment to the Lunar Receiving Laboratory (LRL). The spacecraftwill be offloaded from the ship at Pearl Harbor and transported to an area wheredeactivation of the CM pyrotechnics and propellant system will be accomplished.This operation will be confined to the exterior of the spacecraft. The spacecraft willthen be flown to the LRL and placed in a special room for storage. Contingency planscall for sterilization and early release of the spacecraft if the situation so requires.

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CONTINGENCY OPERATIONSGENERAL

If an anomaly occurs after liftoff that would prevent the space vehicle from followingits nominal flight plan, an abort or an alternate mission will be initiated. Aborts willprovide for an acceptable flight crew and Command Module (CM) recovery whilealternate missions will attempt to maximize the accomplishment of mission obiectivesas well as provide for an acceptable flight crew and CM recovery.

ABORTSThe following sections present the abort procedures and descriptions in order of themission phase in which they could occur.LaunchThere are six launch abort modes. The first three abort modes would result in terminationof the launch sequence and a CM landing in the launch abort area. The remainingthree abort modes are essentially alternate launch procedures and result in insertion ofthe Command/Service Module (CSM) into earth orbit. All of the launch abort modesare the same as those for the Apollo 11 Mission.Earth Parkina OrbitA return to earth abort from earth parking orbit (EPO) will be performed by separatingthe CSM from the remainder of the space vehicle and performing a retrograde ServicePropulsion System (SPS) burn to effect entry. Should the SPS be inoperable, theService Module Reaction Control System (SM RCS) will be used to perform the deorbitburn. After CM/SM separation and entry, the crew will fly a guided entry to a pre-selected target point, if available.Translunar InjectionTranslunar injection (TLI) will be continued to nominal cutoff, whenever possible, inorder for the crew to perform malfunction analysis and determine the necessity of anabort.

Translunar CoastIf ground control and the spacecraft crew determine that an abort situation exists,differential velocity (AV) targeting will be voiced to the crew or an onboard abort pro-gram will be used as required. In most cases, the Lunar Module (LM) will be jettisonedprior to the abort maneuver if a direct return is required. An SPS burn will be


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initiated to achieve a direct return to a landing area. However, a real-time decisioncapability will be exploited as necessary for a direct return or circumlunar trajectoryby use of the several CSM/LM propulsion systems in a docked configuration.Lunar Orbit Insertion

. In the event of an early shutdown of the SPS during lunar orbit insertion (LOI),contingency action w ill depend on the condition which caused the shutdown . If theshutdown was inadvertent, and if specified SPS I imits have not been exceeded, an

. immediate restart wil I be attempted. Upon completion of the LOI burn a real-timedecision will be made on possible alternate missions. If, during the LOI burn, theSPS limits are exceeded, a manual shutdown will be made. The LM Descent PropulsionSystem (DPS) will serve as a backup propulsion system.Descent Orbit InsertionIn the event of a descent orbit insertion (DOI) overburn where, despite trim corrections,the pericynthion remains lower than desired, a so-called bail-out SPS burn will beperformed to raise the pericynthion. Based on Mission Rules criteria, this situationcould lead to a continuation of the nominal landing mission, selection of an alternate

- mission, or early mission termination.Transearth lniectionAn SPS shutdown during transearth injection (TEI) may occur as the result o f aninadvertent automatic shutdown. Manual shutdowns are not recommended. If anautomatic shutdown occurs, an immediate restart will be initiated.ALTERNATE MISSION SUMMARYThe two general categories of alternate missions that can be performed during theApollo 13 Mission are (1) earth orbital and (2) lunar. Both of these categories haveseveral variations which depend upon the nature of the anomaly leading to the alternatemission and the resulting systems status of the LM and CSM. A brief description ofthese alternate missions is contained in the following paragraphs.Earth Orbit

The CSM will dock with the LM, and the photographic equipment will be retrievedfrom the LM. Following this, the LM will be deorbited into the Pacific Ocean areato eliminate debris problems. The CSM will perform SPS plane change maneuvers toachieve an orbital inclination of 40 with daylight coverage of all US passes. Earthorbital photography will then be conducted.

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Lunar OrbitCSM and LM

The nominal mission bootstrap photographic objectives will be accomplished.These objectives include photographs of Censorinus, Descartes, and Davy Rille.The LM will normally be jettisoned prior to accomplishing photographic objectivesto avoid CM window blockage.CSM AloneThe hybrid transfer will be deleted in this case. If the hybrid transfer has beenperformed, the CSM will be placed back on a free return trajectory. A two-burnLOI sequence, as on Apollo 8, 11, and 12 will be used to place the vehicle in a60-NM circular orbit. The LOI b urn will also establish an orbit to pass overCensorinus and Mosting C for photography and landmark tracking.

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MISSION SUPPORTGENERAL

.

Mission Support is provided by the Launch Control Center, the Mission Control Center,the Manned Space Flight Network, and the recovery forces. A comprehensivedescription of the mission support elements is in the MOR Supplement.CONTROL CENTERSThe Launch Control Center (LCC), I ocated at Kennedy Space Center, Florida, is thefocal point for overall direction, control and monitoring of prelaunch checkout,countdown and launch of Apollo/Saturn V Space Vehicles.The Mission Control Center (MCC), I ocated at the Manned Spacecraft Center inHouston, Texas, provides centralized mission control from the time the space vehicleclears the launch tower through astronaut and spacecraft recovery. The MCCfunctions within the framework of a Communications, Command, and Telemetry System(CCATS); Real-T ime Computer Complex (RTCC); Voice Communications System; DisplayControl System; and a Mission Operations Control Room (MOCR) supported by StaffSupport Rooms (SSRs) . These systems allow the flight control personnel to remain incontact with the space vehicle, receive telemetry and operational data which can beprocessed by the CCATS and RTCC for verification of a safe mission, or computealternatives. The MOCR and SSRs are staffed with specialists in all aspects of themission who provide the Flight Director and Mission Director with real-time evaluationof mission progress.MANNED SPACE FLIGHT NETWORKThe Manned Space Flight Network (MSFN) is a worldwide communications and trackingnetwork that is controlled by the MCC during Apollo missions. The network is com-posed of fixed stations supplemented by mobile stations. The functions of these stationsare to provide tracking, telemetry, updata, and voice communications between thespacecraft and the MCC. Connection between these many MSFN stations and theMCC is provided by the NASA Communications Network.

Figure 13 depicts communications during lunar surface operations.Figure 14 shows the MSFN configuration for Apollo 13.

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11-.(D.w

- ,.-..aur\bK VOICE ND TLM)272.5-MHz I PIAVRnr1c2287.5-MHz PRN. VnTrF nm 71s.~106.4-MHzF-W, VOICE,AND UPDATA.

LMP

296.8-Mel: VOICE S-BANDANTENNA

LMLUNARSURFACE OlW1UNICATIONS ETWORK


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RECOVERY SUPPORTGeneralThe Apollo 13 flight crew and Command Module (CM) will be recovered as soon aspossible after landing, while observing the constraints required to maintain biologicalisolation of the flight crew, CM, and materials removed from the CM. After locatingthe CM, first consideration will be given to determining the condition of the astronautsand to providing first-level medical aid if required. The second consideration wi I I berecovery of the astronauts and CM. Retrieval of the CM main parachutes, apex cover,and drogue parachutes, in that order, is highly desirable if feasible and practical.Special clothing, procedures, and the Mobile Quarantine Facility (MQF) will be usedto provide biological isolation of the astronauts and CM. The lunar soil and rocksamples will also be isolated for return to the Manned Spacecraft Center.Primary Landing AreaThe primary landing area, shown in Figure 15, is that area in which the CM will landfollowing circumlunar or lunar orbital trajectories that are targeted to the mid-PacificOcean. The target point will normally be 1250 nautical miles (NM) downrange ofthe entry point (400 ,000 feet altitude). If the entry range is increased to avoid badweather, the area moves along w ith the target point and contains all the high prob-ability landing points as long as the entry range does not exceed 3500 NM.

Figures 16 and 17 show the primary landing area and worldwide recovery forcesdeployment. Recovery equipment and procedures changes fo r Apollo 13 are as follows:l Recovery beacon CM antennas switchedl Sea dye inside CM deployed on requestl Navy type Mae West, astronaut egress lifejacketl New design recovery liferaft

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y$ 3..:

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i i

HAW RES 2ABBREVIATIONS:

R RECOVERY b CHRISTMASP PHOTO

ISLANDS-l SWIM 1ST-l SWIM TEAM 1S-2 SWIM 2ST-2 SWIM TEAM 2HAW RES HAWAIIRESCUE

2?ca.z PRIMARY LANDING AREARECOVERY FORCE VOICE CALLS


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RECOVERY FORCE SUPPORTDESTROYERPEASE I 7-A /- LAJES (1) r-HAWAII (2)\ l-DESTROYER(IN PORT)N 15f 0s 15

- --I ..- -._ .-P I t t -- ,-

-iv-- .i90 45I

0 45 90 135 180 135 9-i-E E-i--WLASCENSION (2)


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CONFIGURATION DIFFERENCESSPACE VEHICLE REMARKSCommand/Service Module (CSM- 109)0 Changed Service Module Jettison

Controller (SMJC) timers.

l Added Lunar Topographic Camerasystem.

0 Added cabin fan filter. Removes lunar dust.Lunar Module (LM-7) (Ascent Stage)0 Added auto and attitude hold

modes in P66 and eliminated P65and P67 software programs.

l Incorporated a non-clogging flowlimiter in primary EnvironmentalControl System lithium hydroxidecanister.

Lunar Module (LM-7) (Descent Stage)0 Installed heat exchanger bypass

I ine on descent stage fuel line.Soacecraft-LM Adapter (SLA-16)l (No significant differences .)

Allows additional SM movement away fromCM prior to RCS termination during sepa-ration. Provides decreased probability ofCM&M recontact during earth entry.

Provides for high resolution lunar topo-graphic photography with image motioncompensation .

Aids crew during lunar landing in obscuredvisibility.

Improves crew comfo rt by eliminatingwater in suit loop.

Facilitates planned depressurization of fueltanks after landing.

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LAUNCH VEHICLE

Instrument Unit (S-IU-508)l Added fourth battery to IU.

l Relocated and added telemetrymeasurements for vibration in-vestigation of ST- 124 inertialplatform.

l Added four wires to IU EmergencyDetection System distributor.

S-m Stage (S-~VB-508)l (No significant differences .)

S-II Stage (S-11-8)

0 lnstal led al I spray foam insulation.

S-IC Stage (S-K-8)l (No significant differences .)

-.

Extends Command Communications Systemtracking to assist S-IVB/IU lunar impacttraiectory corrections.Provides data for analysis if a flight anom-aly occurs on ST- 124.

Provides automatic vehicle ground commandcapability at spacecraft separation in eventof a contingency separation.

Reduces weight of the stage.

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FLIGHT CREWFLIGHT CREW ASSIGNMENTSPrime Crew (Figure 18)

Commander (CDR) - James A. Lovell, Jr. (Captain, USN)Command Module Pilot (CMP) - Thomas K. Mattingly, II

(Lieutenant Commander, USN)Lunar Module Pilot (LMP) - Fred W. Haise, Jr. (Civilian)

Backup CrewCommander (CDR) - John W. Young (Commander, USN)Command Module Pilot (CMP) - John L. Swigert, Jr. (Civilian)Lunar Module Pi lot (LMP) - Charles M. Duke, Jr. (Major, USAF)

The backup crew follows closely the training schedule for the prime crew and func-tions in three significant categories. One, they receive nearly complete missiontraining which becomes a valuable foundation for later assignments as a prime crew.Two, should the prime crew becsome unavailable, the backup crew is prepared to flyas prime crew up until the last few weeks prior to launch. Three, they are fullyinformed assistants who help the prime crew organize the mission and check out thehardware.During the final weeks before launch, the flight hardware and software, groundhardware and software, and flight crew and ground crews work as an integrated teamto perform ground simulations and other tests of the upcoming mission. It is necessarythat the flight crew that will conduct the mission take part in these activities, whichare not repeated for the benefit of the backup crew. To do so would add an additionalcostly and time consuming period to the prelaunch schedule, which for a lunar missionwould require rescheduling for a later lunar launch window.PRIME CREW DATA

CommanderNAME: James A. Lovell, Jr. (Captain, USN)SPACE FLIGHT EXPERIENCE: Captain Lovell was selected as an astronaut by

NASA in September 1962. He has since served as backup pilot for theGemini 4 flight and backup command pilot for the Gemini 9 flight.

-

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On 4 December 1965, he and command pilot Frank Borman were launched onthe Gemini 7 Mission. The flight lasted 339 hours, 35 minutes and includedthe first rendezvous of two manned maneuverable spacecraft as Gemini 7 wasjoined in orbit by Gemini 6.As command pilot, Lovell flew the 4-day, 59-revolution, Gemini 12 Missionin November 1966. This flight, which included a third-revolution rendezvouswith a previously launched Agena, marked the successful completion of theGemini Program.Love1 I served as Command Module Pi lot on the 6-day Apollo 8 (21-27 December1968) first manned flight to the moon. Apollo 8 performed 10 revolutions inlunar orbit and returned to an earth landing after a total flight time of 147 hours.Captain Love1 has since served as the backup spacecraft Commander for Apollo 11,the first manned lunar landing. He has logged a total of 572 hours, 10 minutes ofspace flight in three missions.

Command Module Pi lot- NAME: Thomas K. Mattingly, II (Lieutenant Commander, USN)

SPACE FLIGHT EXPERIENCE: Lieutenant Commander Mattingly was selected as anastronaut by NASA in April 1966. He has served as a member of the astronautsupport crews for the Apollo 8 and 11 Missions.

Lunar Module PilotNAME: Fred W. Haise, Jr. (Civilian)SPACE FLIGHT EXPERIENCE: Mr. Haise was selected as an astronaut by NASA in

April 1966. He has served as a member of the astronaut backup crews for theApollo 8 and 11 Missions.

BACKUP CREW DATACommander

NAME: John. W. Young (Commander, USN)SPACE FLIGHT EXPERIENCE: Commander Young was selected as an astronaut by

NASA in September 1962. He served as pilot on the first manned Geminiflight - a 3-orbit mission launched on 23 March 1965. Following that assign-ment he was backup pilot for Gemini 6.

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On 18 July 1966, Young was the command pilot on the Gemini 10 Missionwhich made two successful rendezvous and dockings with Agena targetvehicles. Gemini 10 was a 3-day, 44-revolution earth orbital flight.He was subsequently assigned as the backup Command Module Pilot forApollo 7.Commander Young flew as the Command Module Pilot on the Apollo 10 lunarorbital mission which performed all but the final minutes of an actual lunarlanding.

Command Module Pi lotNAME: John L. Swigert, Jr. (Civilian)

SPACE FLIGHT EXPERIENCE: Mr. Swigert was selected as an astronaut by NASAin April 1966.

Lunar Module Pilot- NAME: Charles M. Duke, Jr. (Major, USAF)

SPACE FLIGHT EXPERIENCE: Major Duke was selected as an astronaut by NASAin April 1966.

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MISSIONMANAGEMENTRESPONSIBILITYTitle

Director, Apollo ProgramMission DirectorAssistant Mission DirectorDirector, Mission Operations

Saturn Program ManagerMission Operations Manager

Apollo Spacecraft ProgramManagerDirector of Flight OperationsFlight Directors

Spacecraft Commander (Prime)Spacecraft Commander (Backup)

Apollo Program Manager KSCDirector of Launch OperationsLaunch Operations Manager

Name OrganizationDr. Rocco A. Petrone NASA/OMSFCapt. Chester M. Lee (Ret) NASA/OMSFCol. Thomas H. McMullen NASA/OMSFMaj. Gen. John D. Stevenson (Ret) NASA/OMSF

Mr. Roy E. GodfreyDr. Fridtjof A. Speer

Col. JamesA. McDivittMr. Sigurd A. SjobergMr. Milton WindlerMr. Gerald D. GriffinMr. Eugene F. KranzMr. Glynn S. Lunney

NASA/MSFCNASA/MSFC

NASA/MSC

NASA/MSCNASA/MSC

Capt. JamesA. Love11 NASA/MSCCdr. John W. Young NASA/MSCMr. Edward R. Mathews NASA/KSCMr. Walter J. Kapryan NASA/KSCMr. Paul C. Donnelly NASA/KSC

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ABBREVIATIONS AND ACRONYMS

-

AGSALSEPAOSAPSAPSARIAASBIGCCATSCCGECDHCDRCPLEE

ZPcscCSICSMDACDDAS

it:DPSDSKYECSEIEMUEPOESTETBEVAFMfpsFDA1

lilyGNCSGSFCHBRHFEHTCIMUIUIVTKSCLBRLCCLDMKLECLESLETLH4l HZPLO1LOSLOXLPOLRLRLLRRR

Abort Guidance SystemApollo Lunar Surface Experi-ments PackageAcquisition of SignalAscent Propulsi on System (LM)Auxiliary Propulsio n System(S-IVB)Apollo Range InstrumentationAircraftApollo/SaturnBiological Isolation GarmentCmnunications, Ccmrnand, andTelemetry SystemCold Cathode Gauge ExperimentConstant Delta HeightCommanderCharged Particle Lunar Environ-ment ExperimentCommand ModuleCommand Module PilotClose-up Stereo CameraConcentric Sequence InitiationCoimmnd/Service ModuleData Acquisition CameraDigital Data AcquisitionSystemDepartment of DefenseDescent Orbit InsertionDescent Propulsion SystemDisplay and Keyboard AssemblyEnvironmen tal Control SystemEntry InterfaceExtravehicular Mobillty UnitEarth Parking OrbitEastern Standard TimeEquipment Transfer BagExtravehicular ActivityFrequency ModulationFeet Per SecondFlight Director AttitudeIndicatorFixed Throttle PositionGround Elapsed TimeGuidance, Navigation, andControl System (CSM)Goddard Space Flight CenterHigh Bit RateHeat Flow ExperimentHand Tool CarrierInertial Measurement UnitInstrument UnitIntravehicular TransferKennedy Space CenterLow Bit RateLaunch Control CenterLangarkLunar Equipment ConveyorLaunch Escape SystemLaunch Escape TowerLiquid HydrogenLithium HydroxideLunar ModuleLunar Module. PilotLunar Orbit InsertionLoss of SignalLiquid OxygenLunar Parking OrbitLanding RadarLunar Receiving LaboratoryLaser Ranging Retro-Reflector

LSMLVMCCMCCMESAMHZMOCRMORMPLWMSCMSFCMSFNNASCOM:;SFOPSORDEALPCMPDIPGAPGNCSPLSSPSEPTCQUA0RCSRRRLSRTCCRTGSICSEAs-ICS-IIS-IVBSIDE

Lunar Surface MagnetometerLaunch VehicleMidcourse CorrectionMission Control CenterModularized Equipment StowageAssemblyMegahertzMission Operations Control RoanMission Operations ReportMid-Pacific LineMobile Quarantine FacilityManned Spacecraft CenterMarshall Space Flight CenterManned Space Flight NetworkNASA Camnunica tions NetworkNautical MileOffice of Manned Space FlightOxygen Purge SystemOrbital Rate Display Earth andLunarPulse Code ModulationPowered Descent InitiationPressure Garment AssemblyPrimary Guidance, Navi(LM3

ation,and Control SystemPortable Life Suppor t SystemPassive Seismic ExperimentPassive Thermal ControlOuadrant--m.-Reaction Control SystemRendezvous RadarRadius Landing SiteReal-Time Computer CanplexRadioisotope ThermoelectricGeneratorSpacecraftSun Elevation Anole

SLA94SPANSPSSRCSSBSSRsvSWCTD6ETECTEITFITLCTLITLMTPFibiT-timeTVUSEUSNUSAFVANVHFAV

Saturn V First StageSaturn V Second StageSaturn V Third StageSuprathennal Ion DetectorExperimentSpacecraft-LM AdapterService ModuleSolar Particle Alert NetworkService Propulsi on SystemSample Return ContainerSingle Side BandStaff Support RoomSpace VehicleSolar Wind CaapositionExperimentTransposition, Docking and LMEjectionTransearth CoastTransearth InjectionTime From IgnitionTranslunar CoastTranslunar InjectionTelemetryTerminal Phase Finalizati onTerminal Phase InitiationCountdow n Time (referenced toliftoff time)TelevisionUnified S-BandUnited States NavyUnited States Air ForceVanguardVery High FrequencyDifferential Velocity_.-

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Post LaunchMission Operation ReportNo. M-932-70-1 3

TO: A/Ad ministrator 28 April 1970FROM: MA/ApoI lo Program DirectorSUBJECT: Lip01 IO 13 Mission (AS-508) Post L aunch Mission Operation Report No. 1

The Apollo 13 Mission was successfully launched from Kennedy Space Center, Floridaon Saturday, 11 April 1970. Apollo 13 was progressing smoothly to a planned lunarlanding until about 56 hours into the flight when a failure occurred in the ServiceModule cryogenic oxygen system. This resulted in a loss of capability to generateelectrical power, to provide oxygen, and to produce water in the Command/ServiceModule. The decision was made to not perform the lunar landing mission and to returnto earth using the Lunar Module for life support, power, propulsion, and guidance.Safe recovery of the crew and Command Module took place in the Pacific Oceanrecovery area on Friday, 17 Apri I 1970. An intensive investigation has been initiatedto determine the cause o f the anomaly.

The Mission Directors Summary Report for Apollo 13 is attached and submitted as PostLaunch Mission Operation Report No. 1. Also attached are the NASA OMSF PrimaryMission Objectives for Apollo 13. Since these Primary Objectives could not beachieved without a lunar landing, I am recommending that the Apollo 13 Mission beconsidered unsuccessful. Detailed analysis of all data will continue and appropriaterefined results of the mission wil I be reported in the Manned Space Flight Centerstechnical reports.

kfLM=-Rocco A. PetroneAPPROVAL:


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PRIMARY OBJECTIVES

M-932-70- 13

NASA OMSF PRIMARY MISSION OBJECTIVES FOR APOLLO 13

. Perform selenological inspection, survey, and samplpreselected region of the Fra Mauro Formation.

ing of materials in a

. Deploy and activate an Apollo Lunar Surface Experi ments Package (ALSEP).

. Develop mans capability to work in the lunar environment.

. Obtain photographs of candidate exploration sites.

Apol lo Program DirectorManned Space Flight

RESULTS OF APOLLO 13 MISSION

Apollo 13, launched 11 Apri I 1970, was aborted after 56 hours of flight and terminatedon 17 April 1970. The planned lunar landing was not accomplished and this mission isadiudged unsuccessfu l in accordance with the objectives stated. above.

ELS.BZURocco A. PetroneApollo Program Director

Manned Space FlightA7 r;f. /470/I

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NATIONAL AERONAUTICS AND SPACE ADMINISTRATIONWASHINGTO N, D.C. 20546

TO: DistributionFROM: MA/ApoI IO Mission Director

20 April 1970

SUBJECT: Mission D Ni ectors Summary Report, Apollo 13

INTRODUCTIONThe Apollo 13 Mission was planned as a lunar landing mission but was aborted enrouteto the moon during the third day of flight due to loss of Service Module cryogenicoxygen and consequent loss of capability to generate electrical power, to provideoxygen, and to produce water in the Command/Service Module. Shortly after theanomaly, the Command/Service Module was powered down and the remaining flight,except for entry, was made with the Lunar Module providing necessary power, environ-mental control, guidance, and propulsion. Flight crew members were: Commander(CDR), Capt. J ames Lovell, Jr.; Command Module Pilot (CMP), Mr. John Swigert, Jr.;Lunar Module Pilot (LMP), Mr. Fred W. Haise, Jr, Swigert, officially the backupCMP for the Apollo 13 Mission, was substituted for LCDR Thomas K. Mattingly, II,the prime crew CMP, when it was feared that Mattingly had possibly contractedRubella, and if so, could be adversely affected in performing his demanding auties.A vigorous simulation program was successfully completed prior to launch to ensurethat Lovell, Swigert, and Haise could function with unquestioned teamwork througheven the most arduous and time-critical simulated emergency conditions. Significantdetailed mission data are contained in Tables 1 through 4.PRELAUNCHNo problems occurred during space vehicle prelaunch operations to impact the count-down. However, the S-IC Stage No. 2 liquid oxygen (LOX) vent valve did not closewhen commanded at T minus 1 hour 58 minutes. After cycling the valve several timesand flowing ambient nitrogen gas through the valve, it was successfully closed at T minus1 hour 21 minutes. Weather conditions at launch were: overcast at 20,000 feet,visibility 10 miles, wind 10 knots.


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LAUNCH AND EARTH PARKING ORBIT

Apollo 13 was successfully launched on schedule from Launch Complex 39A, KennedySpace Center, Florida, at 2:13 p.m. EST, 11 April 1970. The launch vehicle stagesinserted the S-IVB/Instrument Unit (IU)/ p acecraft combination into an earth parkingorbit with an apogee of 100.2 nautical miles (NM) and a perigee of 98.0 NM (loo-NMcircular planned). During second stage boost the center engine of the S-II Stage cut offabout 132 seconds early causing the remaining four engines to burn approximately 34seconds longer than predicted. Space vehicle velocity after S-II boost was 223 feet persecond (fps) lower than planned. As a result, the S-IVB orbital insertion burn wasapproximately 9 seconds longer than predicted with cutoff velocity within about 1 .2 fpsof planned. Total launch vehicle burn time was about 44 seconds longer than predicted.A greater than 3-sigma probability of meeting translunar injection (TLI) cutoff conditionsexisted with remaining S-IVB propellants.

After orbital insertion, all launch vehicle and spacecraft systems were verified andpreparations were made for TLI . Onboard television was initiated at 01:35 GET (hour:minutes ground elapsed time) for about 5.5 minutes. The second S-IVB burn wasinitiated on schedule for TLI. All major systems operated satisfactorily and all endconditions were nominal for a free-return circumlunar trajectory.TRANSLUNAR COASTThe Command/Service Module (CSM) separated from the Lunar Module (LM)/IU/S-IVBat about 03:07GET. Onboard television was then initiated for about 72 minutes andclearly showed CSM hard docking, ejection of the CSM/LM from the S-lVB/IU atabout 04:Ol GET, and the S-IVB Auxiliary Propulsion System (APS) evasive maneuveras well as spacecraft interior and exterior scenes. Service Module Reaction ControlSystem (SM RCS) propellant usage for the separation, transposition, docking, andejection was nominal. All launch vehicle safing activities were performed as scheduled.The S-IVB APS evasive maneuver by an 8-second APS ullage burn was initiated at04:18 GET and was successfully completed. The LOX dump was initiated at 04:39 GETand was also successfully accomplished. The first S-IVB APS burn for lunar target pointimpact was initiated at 06:OO GET. The burn duration was 217 seconds producing adifferential velocity of approximately 28 fps. Tracking information available at 08:OOGET indicated that the S-lVB/IU would impact at 653S, 3053W versus the targeted35, 3OW. Therefore, the second S-IVB APS (trim) burn was not required. The gaseousnitrogen pressure dropped in the IU ST-124-M3 inertial platform at 18:25 GET and theS-lVB/IU no longer had attitude control but began tumbling slowly. At approximately19:17 GET, a step input in tracking data indicated a velocity increase of approximately4 to 5 fps. No conclusions have been reached on the reason for this increase. Thevelocity change altered the lunar impact point closer to the target. The S-lVB/IUimpacted the lunar surface at 77:56:40 GET (08:09:40 p.m. EST, 14 April) at 2.4S,

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27.9W and the seismometer deployed during the Apollo 12 Mission successfully detectedthe impact (see MISSION SCIENCE). The targeted impact point was 125 NM fromthe seismometer-. The actual impact point was 74 NM from the seismometer, well withinthe desired 189-NM (350 -kilometer) radius.

The accur-acy of the TLI maneuver was such that spacecraft midcourse correction NO. 1(MCC- l), scheduled for 11:4 1 GET, was not required. MCC-2 was performed as plannedat 30:41 GET and resulted in placing the spacecraft on the desired, non-free-retu rncircumlunar trajectory with a predicted closest approach to the moon of 62 NM. AllService Propulsion System (SPS) burn parameters were normal. The accuracy of MCC-2was such that MCC-3, scheduled for 55:26 GET, was not performed. Good qualitytelevision coverage of the preparations and performance of MCC-2 was received for49 minutes beginning at 30:13 GET.

At approximately 55:55 GET (lo:08 p.m. EST) the crew reported an undervoltage alarmon the CSM Main Bus B. Pressure was rapidly lost in Service Module oxygen tankNo. 2 and fuel cells 1 and 3 current dropped to zero due to loss of their oxygen supply.A decision was made to abort the mission. The increased load on fuel cell 2 anddecaying pressure in the remaining oxygen tank led to the decision to activate theLM, power down the CSM, and use the LM systems for life support.At 61:30 GET, a 38-fps midcourse maneuver (MCC-4) was performed by the LM DescentPropulsion System (DPS) to place the spacecraft in a free-return trajectory on which theCommand Module (CM) would nominally land in the Indian Ocean south of Mauritius atapproximately 152:00 GET.TRANSEARTH COAST_---At pericynthion plus 2 hours (79:28 GET), a LM DPS maneuver was performed to shortenthe return trip time and move the earth landing point. The 263.4-second burn produceda differential velocity of 860.5 fps and resulted in an initial predicted earth landingpoint in the mid-Pacific Ocean at 142:53 GET. Both LM guidance systems were poweredup and the pt-imary system was used for this maneuver. Following the maneuver, passivethermal control was established and the LM was powered down to conserve consumables;only the LM Envit-onmental Control System (ECS) and communications and telemetrysystems were kept powered up.The LM DPS was used to perform MCC-5 at l&5:19 GET. The 15-second burn (at 10%throttle) produced a velocity change of about 7.8 fps and successfully raised the entryflight path angle to -6.52.The CSM was partially powered up for a check of the thermal conditions of the CM withfirst reported receipt of S-band signal at lO1:53 GET. Thermal conditions on all CSMsystems observed appeared to be in order for entry.

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Due to the unusual spacecraft configuration, new procedures leading to entry weredeveloped and verified in ground-based simulations. The resulting timeline called fora final midcourse correction (MCC-7) at entry interface (El) -5 hours, jettison of theSM at El -4.5 hours, then jettison of the LM at El -1 hour prior to a normal atmosphericentry by the CM.MCC-7 was successfully accomplished at 137:40 GET. The 22.4-second LM RCSmaneuver resulted in a predicted entry flight path angle of -6.49. The SM wasjettisoned at 138:02 GET. The crew viewed and photographed the SM and reportedthat an entire panel was missing near the S-band high-gain antenna and a great dealof debris was hanging out. The CM was powered up and then the LM was jettisonedat 141:30 GET. The El at 400,000 feet was reached at 142:41 GET.ENTRY AND RECOVERY

Weather in the prime recovery area was as follows: Broken stratus clouds at 2000 feet;visibility 10 miles; &knot ENE winds; and wave height I to 2 feet. Grogue and mainparachutes deployed normally . Visual contact with the spacecraft was reported at142:50 GET. Landing occurred at 142:54 :41 GET (01:07:41 p.m. EST, I7 April). The

ately 2 l405, 165O22W. Theanding point was in the mid-Pacific Ocean, approximCM landed in the stable 1 position about 3.5 NM fro mIWO JIMA. The crew was picked up by a recovery heship at 1:53 p.m. EST, less than an hour after landing.

the-prime recovery ship, USSlicopter .and was safe aboard the

MISSION SCIENCEThe S-IVB Stage, weighing about 30,700 pounds, impacted the moon 74 NM from theApollo 12 seismometer at an angle of about 80 to the horizontal with a velocity of8465 fps and an energy equivalency of 11.5 tons of TNT. These data compare with theApollo 12 LM, which hit the moon at a distance of 42 NM from the seismometer at anangle of 3 to the horizontal, and an equivalent energy of approximately 1 ton of TNT.The overall character of the seismic signal is similar to that of the LM impact signal,but the higher impact energy gave a seismic signal 20-30 times larger than the LMimpact and 4 times longer in duration (approximately 4 hours vs I hour). The signalwas so large that the gain of the seismometer had to be reduced by ground command tokeep the recording on scale. A clear signal was recorded on the three long periodcomponents so that it is possible to distinguish each event with absolute certainty.Thirty seconds elapsed between time of impact and arrival of the seismic wave at theseismometer; peak amplitude occurred 7 minutes later.The signal arrival time had been predicted on the basis of velocity measurements madeon the Apollo 11 and I2 lunar sample materials in the laboratory. The average velocityof the seismic wave through the lunar material is 4.6 km/set which compares favorably


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with the 3.>km/sec velocity recorded by the LM impact. The depth of penetration ofthe S-IVB impact signal is believed to be 20-40 km (vs 20 km for LM impact). Thisresult implies that the outer shell o f the moon, to depths of at least 20-40 km, may beformed of the same crystalline rock material as found at the surface. No evidence ofa lower boundary to this material has been found in the seismic signal, although it isclear that it is too dense to form the entire moon.One puzzling feature of the signal is the unexpectedly rapid build-up from the beginningto its maximum. This part of the signal, at least, cannot be satisfactorily explained by

,scattering of seismic waves in a rubble material as was thought possible from the earlierLM impact data. Scattering of signals may explain the later part of the signal. Severalalternate hypotheses are under study, but no firm conclusions have been reached. Onepossibility is that the expanding cloud of material from the impact produces seismicsignals continuously as it sweeps across the lunar surface.The fact that such precise targeting accuracy was achieved for the S-IVB and that theresulting seismic signals were so large have greatly encouraged scientists to believethat planned future impacts can be extended to ranges of at least 500 km and that thedata return w ill provide the means for determining the structure of the moon to depthsapproaching 200 km.The Suprathermal Ion Detector Experiment (SIDE), also part of the ALSEP 1 experimentspackage deployed during the Apollo 12 Mission, recorded a jump in the number of ioncounts a fter the S-IVB impact. Since the instrument was in lunar shadow at the time ofimpact the ion count was essentially zero. A few ions were recorded 22 seconds afterimpact; a second frame of data showed a iump to 250 ions, the third jumped to 2500ions, the fourth dropped back to a few ions, then the count fell back essentially tozero. These ions were in the 70 electron voltenergy range. All of the counts wereobserved over a period of 70 seconds. In addition to the ion counts, the mass analyzerof the instrument also recorded ions, almost all of which were in the 50-80 mass unitrange (hydrogen = 1 mass unit).Two possible mechanisms have been given for producing ions: (1) temperatures in theranges 6000-10 ,OOOC generated by the S-IVB impact could produce ionization; (2)particles that reach heights of 60 km could also be ionized by sunlight.SYSTEMS PERFORMANCESaturn V S-IC ignition, holddown arm release, and liftoff were accomplished withinexpected limits and indications are that S-IC systems performed at or near nominal.LOX tank pressure was as expected throughout the burn.

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All S-II Stage systems were nominal throughout S-II burn until the center J-2 engineshut down approximately 132 seconds earlier than scheduled. LOW frequency oscillations(14 to 16 h -t )r z ex p erienced on the S-II Stage resulted in a 132-second premature centerengine cutoff. Preliminary analysis indicates that a Thrust OK pressure switch cutoffoccurred due to large pressure oscillations in the LOX system. No apparent engine orstructural damage was incurred. Oscillations in the stage and outboard engines decayedto a normal level follow ing center engine cutoff. Preliminary data does not indicateany off-nominal performance of the four outboard engines.All S-IVB systems operated within expected limits during both the first and second burns.The first burn was 9.2 seconds longer than predicted, making up for the velocity deficitat S-II cutoff. The second (TLI) b urn was approximately 5 seconds longer than predictedfrom observed orbital conditions. A small vibration was reported by the crew approxi-mately 90 seconds prior to second burn cutoff.All IU guidance and control functions were satisfactory and all systems performed asexpected.Performance of the CSM fuel cell and cryogenic systems was nominal until 55:53:36 GETwhen an unusual pressure rise was noted in oxygen tank No. 2. The pressure continued torise to the relief valve crack pressure of 1004.1 psia (pounds per square inch absolute).One second later,at 55:54:45 GET, the pressure reacheda maximum of 1008.3 psia at whichtime the tank vent valve apparently opened. The last valid tank pressure reading prior toloss of data was 995.7 psia at 55:54:53 GET. At 55:54:54 GET an undervoltage cautionlight occurred on Main Bus B, which was powered by fuel cell 3. Concurrent with theabrupt loss of oxygen tank No. 2 pressure, oxygen tank No. 1 pressure showed a rapiddecrease to about 373 psia in 87 seconds. Fuel cells 1 and 3 were removed from the lineabout 18 minutes after the anomaly. Fuel cell 2 remained in operation for about 2 hoursbefore the oxygen pressure in tank No. 1 had decreased to 61 psia and the fuel cell wasremoved from the line. As a result of these occurrences, the CM was powered down andthe LM was configured to supply the necessary power and other consumables.

Power down of the CSM began at 58:40 GET. Th e surge tank and repressurizationpackage were isolated with approximately 860 psi residual pressure (approximately 6.5pounds of oxygen total). The primary water glycol system was left with radiators by-passed. Indicated water tank residuals were 18.0 pounds in the waste tank and 37.5pounds in the potable tank. All SM RCS quads were powered down with heatersdeactivated. All SPS parameters were nominal before and after the anomaly and noconfiguration changes took place after the anomaly.All LM systems performed satisfactorily in providing the necessary power and environ-mental contro l to the spacecraft. The requirement for lithium hydroxide to removecarbon dioxide from the spacecraft atmosphere was met by a combination of CM andLM cartridges since the LM cartridges alone wo uld not satisfy the total requ irement.

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The crewmen, with instructions from Mission Control, built an adapter from the CMcartridges to accept the LM hoses.The LM supercritical helium (SHe) tank pressure exhibited an increased rise rate afterthe second DPS firing. Prior to the burn, the rise rate had been 11 psi per hour. Afterthe burn, the rate increased to 33 psi per hour. After the third DPS maneuver, the SHetank burst disc ruptured at 108:54 GET at a pressure of about 1940 psi, within theexpected range. The passive thermal control mode in use at the time was affected bya small attitude rate change from the venting SHe changing from a right yaw rate of0.3O/sec to a left yaw rate of 3.0sec, but did not cause any problem.The CSM was partially powered up at about 101:53 GET with the following results:

Telemetry - Following a brief period of intermittent S-band reception, solidtelemetry was received from 101:49 GET to system power-down at 102:03 GET.Telemetry system performance was nominal throughout the time period it waspowered up.Instrumentation - A summary review indicated no discrepancies. The centraltiming equipment updated correctly in resetting to 0 (zero) and indicatingaccumulated time from the turn-on associated with the status check of the CSMat 101:53 GET.Electrical Power - All system bus voltage and inverter performance was nominal.Only Main Bus B, Battery Bus B, and AC Bus 1 were powered up. Prior toinstrumentation power-up, the three entry batteries had been on true opencircuit (i.e., no parasitic loads) since approximately 58:40 GET. All performanceto that point had been nominal. CSM Main Bus B was powered up using Battery Band performance under load was nominal. Approximately 2.5 ampere-hours wereconsumed. Battery A, which was used to supplement CM power immediatelyfollowing the fuel cell anomaly, was recharged from the LM ascent batteries.Battery

apollo 13 mission operation report

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