ALTERNATE LUNAR MISS

APOLLO lQ'(MISSION F)

SPACECRAFTiOPERATIONAl

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Lunar Mission Analysis Branch

M!SSION PLANNING AND ANAl'SiS DIVISION

MANNED SPACECRAFT CENTERHOUSTON.TEXAS

Unclas00/99 16320

}~:~:~{{~~ASA-TM-X-69706) APOLLO 10 (MISSION F):::::::::;:::::, SPACECRAFT OP ERATIONAL ALTERN AT EMISSION::::::::::::::::PLANS. VOLUME 2: ALTEENATE LUNAR::::::::::::::::"ISSIONS (NASA) 42 p

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MSC INTERNAL NOTE NO. 69-FM-87

PROJECT APOLLO

APOLLO 10 (MISSION F) SPACECRAFTOPERATIONAL ALTERNATE MISSION PLANSVOLUME II - ALTERNATE LUNAR MISSIONS

By Rocky D. DuncanLunar Mission Analysis Branch

April 14, 196.9

MISSION PLANNING AND ANALYSIS DIVISION

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

MANNED SPACECRAFT CENTER

HOUSTON, TEXAS

Approved:~ <".~Ronan::serry;Chief ~Lunar Mission Analysis Branch

AP:roved:M..r&~~ Mission Planning and Analysis Division

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• Section

CONTENTS

Page

1.1.1 Nonnominal TLI requirement for an MCCwhich would preclude the nominalmission profile•

1.0 SUMMARY AND INTRODUCTION

1.1 Alternate 1 ... ..1

1

1.2 Alternate 2 . . . . . 2

•2.0

3.0

1.2.1 Failure to perform T, D, and E

1.3 Alternate 3 ..

1.3.1 LM NO-GO for undocking and rendezvous

u 1.4 Alternate 4 ..

1.4.1 CSM communications failure in lunarorbit

ABBREVIATIONS

DISCUSSION . .

3.1 Alternate 1 - Contingency: Nonnominal TLIRequirement for MCC Which Would Preclude theNominal Mission Profile . . . . . . . . .

2

2

2

2

2

3

5

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3.1.1 Alternate la - CSM/LM lunar orbitalmission with DPS LOI . . . . . . . 5

3.1.2 Alternate Ib - LM testing during trans-lunar coast with CSM-only lunarorbital mission . . . . . 10

3.1.3 Alternate lc - CSM/LM f~yby 10

Undoeking and Rendezvous

3.2 Alternate 2: CSM-only Lunar Orbital Mission

3.3 Alternate 3

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3.3.13.3.2

Contingency: LM NO-GO for

Alternate 3a - DPS TElAlternate 3b - LM APS burn to depletion

in lunar orbit . . . . . . . . . . . .

iii

11

13

14

Section

3. 1j Alternate 4 - SPS TEl with Docked APS Stage

REFERENCES . •

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Page

14

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•FIGURES

Figure Page

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2

3

4

5

6

8

9

10

11

12

13

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15

eSM-only ~V remaining following a eSM/LM SPS burn .

Mean TEl ~V requirements versus TE flight time forMay 1969 launch window . . . . •

~v resulting from a 40-second SPS burn versus massat burn i~tiation. . . . .

Flow of service propulsion propellants

DPS ~V capability based on low fuel loadings

TLI burn profile for May 17, i969, 72° launch azimuthfirst. injection opportunity .

Mee ~v requirements for various apogee ellipsesresulting from premature S-IVB shutdowns

Propellant used and resultant mass following adocked SPS burn . . . . • . . . . . . . .

Docked DPS ~v capability for various masses atburn initiation . .

Nominal SPS LOI for 72° !lunch azimuth first op­portunity for May 17 launch window

DPS LOI into a 60- by 707-n. mi. orbit

TLI + 25 hours APS burn to depletion

Events sequence for LOI day

Event sequence for Dar day

SPS DOl . . . . .

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16

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25

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16 Landmark geometry for 8- by 60-n. mi. lunar orbit

(a) Perilune 16° behind landmark ..(b) Elevation versus time for perilune of

60- by 8-n. mi. orbit .... .(c) Inertial altitude for landmark tracking a

perilune of 60- by 8-n. mi. orbit

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31

32

Figure

17 DPS TEl on DOl day • • . . • • • . • • • • • • • . • •

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flage

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APOLLO 10 (MISSION F) SPACECRAFT

OPERATIONAL ALTERNATE MISSION PLANS

VOLUME II - ALTERNATE LUNAR MISSIONS

By Rocky D. Duncan

1.0 SUMMARY AND INTRODUCTION

This document is the second of three" volumes entitled Apollo 10(Mission F) Spacecraft Operational Alternate Mission Plans. Specifically,this document presents the alternate lunar missions for Apollo 10 (Mis­sion F) exclusive of the lunar rendezvous plans.

The guidelines used for the lunar alternate missions were thefollowing.

1. Priority of LM testing over CSM testing

2. Operations within the lunar mission time line as closely aspossible

3. No requirement for additional crew training

The alternate missions presented in this document are summarized inthe following paragraphs. Decision logic for the alternate lunar mis­sions is presented in flow chart 1.

1.1 Alternate 1

1.1.1 Nonnominal TLI requirement for an MCC which would precludethe nominal mission profile.-

1. Alternate la - CSM/LM lunar orbital mission with DPS LOI

2. Alternate Ib - LM testing during translunar coast with CSM-onlylunar orbital mission

3. Alternate Ie - CSM/L~ flyby

2

1.2 Alternate 2

1.2.1 Failure to perform T, D, and E.- Alternate 2 is a CSM-onlylunar orbital mission.

1.3 Alternate 3

1.3.1 LM NO-GO for undocking and rendezvous.-

1. Alternate 3a - Docked DPS TEl

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2. Alternate 3b APS burn to depletion in lunar orbit.

1.4 Alternate 4

1.4.1 CSM communications failure in lunar orbit.- Alternate 4 isan BPS TEl with APS stage.

Of prime consideration for the alternate missions presented in thisdocument are the crossovers between the use of one alternate instead ofanother. For example, what is the deciding factor between a CSM-onlylunar orbital mission and a flyby mission? The decision depends on theend of mission ~V reserve philosophy. This philosophy was never resoivedfor Apollo 8 and prior to publication of this document had not been re­solved for Apollo 10 (Mission F). It is the purpose of this documentnot only to present the alternates but also to present data so that thedecision logic between alternates can be resolved for this mission.

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APS

CDR

CLA

CSI

CSM

DOl

DPS

EI

FCSD

FTP

g.e.t.

G&N

I sp

1M

LOI

LOS

LPO

MCC

MSC

NAR

PC

RCS

SPS

3

2.0 ABBREVIATIONS

ascent propulsion system

constant differential height

contingency landing area

coelliptic sequence initiation

command and service modules

descent orbit insertion

. descent propulsion system

entry interface

Flight Crew Support Division

full throttle position

ground elapsed time

guidance and navigation

specific impulse

lunar module'

lunar orbit insertion

loss of signal

lunar parking orbit

midcourse correction

Manned Spacecraft Center

North American Rockwell

pericynthion

reaction control system

service propulsion system

T ~ D~ andE

TEC

TEl

TLC

TLl

TPl

t-.V

4

transposition~ docking~ and extraction

transearth coast

transearth injection

trans lunar coast

trans lunar insertion

terminal phase initialization

change in velocity

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3.0 DISCUSSION

3.1 Alternate 1 - Contingency: Nonnominal TLI Requirementfor MCC Which Would Preclude the Nominal Mission Profile

Alternates la~ lb, and lc are concerned with the nonnominal TLIsituations from which it would not be possible to do an MCC to returnto the nominal mission. In this case~ T, D, and E would be performed,as in the nominal time line. After T~ D~ and E had been performed~ adocked SBS maneuver of up to 4000 fps would be made to place the CSM/1Mon a free-return circumlunar trajectory. Based on a burn of this magni­tude, the resultant CSM/LM weight configuration would result in a DPSabort capability of approximately 3000 fps. An additional 3300 fpsabort capability is available by jettison of the 1M.

3.1.1 Alternate la - CSM/LM lunar orbital mission with DPS LOI.­Alternate lawould be used for a situation in which the first SPS MCCwould preclude an SPS, LOI-l, LOI-2,.and TEl.

The decision to perform an SPS LOI would be a function of the endof mission ~V reserve philosophy.

The following ~V requirements were taken from reference 1 for

the May 1969 launch window.~

1. Range of LOI-l, ~V ~ 2950 fps to 3040 fps

2. LOI-2 6V ~ 138 fps

3. Range of TEl 6V ~ 2658 fps to 4150 fps

The TEl 6V requirements are based on transearth flight times fromapproximately 41 to 126 hours. Other considerations in the reserve

bbudget are as follows.

1. CSM rescue in lunar orbit = 190 fps

2. Translunar MCC 6V = 120 fps

~e nominal trans earth flight time for the May 18 window rangesfrom 41 to 53 hours. This time is not reflected in reference 2.

b These would not be totaled since this would be allowing for doublefailure or contingencies (i.e., CSM rescue and weather avoidancerequirements).

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3. Weather avoidance at EI minus 24 hours = 800 fps

4. G&~ failure in lunar orbit ~ 120 fps

If a large SPS MCC were required because of a nonnominal TLI, thenthe items listed on the previous page would have to be traded off beforecommitment to an SPS 1,01. The major consideration in the f::.V reservephilosophy is the trans earth flight time requirement. The other absolute6V requirements (LOI-l and LOI-2) are relatively inflexible.

The CSM-only 6V capability following any docked SPS burn is presentedin figure 1, and f::.V requirements compared to transearth flight time across

the May 1969 launch window are presented in figure 2. a If the docked SPSburns (MCC, L01-1, and LOI-2) are summed and if the sums are used withfigures 1 and 2, the transearth flight times that can be achieved can bedefined. Because the discrete solutions to the 1650 W CLA for eachlaunch day also are shown in figure 2, the approximate end of mission6V reserves can be derived based on a return to this landing site. Theend of mission 6V reserve philosophy was not resolved for Apollo 8 inthe mission rules and has not yet been resolved for Apollo 10 (Mission F).

If for anyone of the considerations listed above it were decidednot to do an SPS LOI, then a docked DPS LOI could be performed. TheDPS LOI profile would be to burn the DPS to insert the SC into a highapocynthion lunar orbit. Subsequently, the SPS would be used to cir­cularize the orbit to the nominal 60-n. mi. orbit. The following con­straints and guidelines apply to using the DPS for LOI.

1. The L01-1 maneuver should be targeted such that the subsequentSPS burn is a minimum of 40 seconds in duration. This condition is basedon the constraint in reference 2 which states that after a docked SPSburn, the next BPS burn must be at least 40 seconds in duration toguarantee SPS multirestart capability. The constraint will insure thatany helium will be cleared out which might have accumulated in the SPSfeed lines because of the large negative ullage. The CSM-only 6V thatresults from a 40-second burn is compared to mass at burn initiationin figure 3.

aA mean curve of transearth flight time compared with f::.V for theMay launch window is presented in figure 2. The curve is in error upto 50 fps for some days. The data have been generatpd in detail acrosseach daily launch window and will be documented in an ~1SC internal note.

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I2. In addition to the first constraint, NR has further recommended~

that because of the criticality of the SPS in lunar orbit the DPS LOIshould not be performed unless the SPS sump tanks are full. The storageand sump tank configurations are generally illustrated in figure 4. Be­cause the storage tanks are depleted first and they contaim about one-halfof the SPS fuel, for the sump tanks to be full at LOI requires that nomore than one-half of the SPS fuel can be burned prior to LOI. Based onF mission loading, half of the SPS fuel is 20 108 pounds, which correspondsto a docked SPS ~V of approximately 2400 fps prior to LOI.

,3. The DPS should not be burned to depletion. Enough fuel shouldbe held in reserve to perform the nominal mission rendezvous profile.The rendezvous ~v requirements are as follows.

Burn ~V, fps Propulsion system

DOl 73 DPS

Phasing 193 DPS

Insertion 213 APS

CSI 50 RCS

CDR 6 RCS

TPI 25 RCS

Total 560

It is recommended that 1000 pounds of DPS fuel be reserv~d forrendezvous. This allotment is quite conservative because based on evena low DPS I of 300 seconds it ~ill yield apprOXimately 685 fps, whichsp .is more than sufficient to fly the entire rendezvous profile with the DPSif required. If a smaller reserve were considered, the DPS ~v capa­bility based on a low fuel loading (0 to 1200 Ib) can be determined fromfigure 5.

Apogee altitude compared to burn time for a typical TLI burn ispresented in figure 6. Because apogee altitude increases so rapidlynear the end of the burn, a very small premature shutdown can result ina large MCC to return to the nominal !~·ajecto~y. Typical MCC ~v

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requirements at 2, 3, 5, and 7 hours after TLI cutoff are shown infigure 7 for various apogee ellipses that result from premature S-IVB

shutdowns during TLI. a

The SPS fuel used and the mass that results after docked SPS burnare presented in figure 8. Finally, docked DPS ~V capability based onCSM/LM mass at burn initiation is presented in figul'e 9. The resultantDPS 6V capability for LOI can be ascertained from these curves, basedon SPS propellant required to correct a nonnominal TLI.

All the curves are presented so that the ~V tradeoffs involvedin Commitment to a DPS LOI can be shown~ This commitment would involvesome very complicated decision logic if it were not for constraint2 listed above; that is, LOI will not be performed with the DPS unlessthe SPS sump tanks are full. The constraint dictates that no more thanhalf of the SPS fuel be used prior to LOI (20 108 Ib) and that thetotal pre-LOr SPS ~V be less than 2421 fps. The resultant CSM/LM masswould produce a docked DPS ~V capability of 2460 fps with 1000 poundsof propellant reserved for rendezvous.

The ~V and resultant apogee altitudes are presented in figure 10as a function of burn time for an LOI burn on the May 17 launch window.The previously described DPS ~V would insert the CSM/1M into approximatelya 60- by 700-n. mi. orbit (fig. 10). It also can be seen from figure 10that even with the minimum DPS ~V of 1950 fps (full SPS tanks) the LMwould be capable of insertion into a 60- by 1600-n. mi. orbit. Therefore,based on the previously listed ground rules the range of lunar orbitsof which the DPS ~s capable are from 60 by 700 n. mi. to 60 by 1600 n. mi.

The cutoff point for the decision to perform an SPS burn or a LMDPS LOI will depend on the end of mission ~V reserve philosophy.

For this document, a DPS LOI for the May 17 launch window wassimulated based on the following profile.

1. The first MCC used half of the SPS fuel (~V = 2421 fps).

2. The CSM/1M mass at LOr was 74 488 pounds.

3. The DPS lunar ~V was near the maximum capability while a1000-pound fuel reserve was maintained for rendezvous.

aThese data were based on a premature S-IVB shutdown and do not

consider any guidance dispersions. The data are provided as typicaldata and can be expected to vary somewhat for different launch azimuthsand launch days.

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4. The burn profile was 10 percent thrust for 15 seconds, and theremainder of the burn was performed at FTP.

A summary of the DPS LOI burn is presented in the following table .

Time of LOI initiation, hr:min:sec, g.e.t~

Selenographic longitude of initiation., deg W .•

5. The average I was assumed to be 302.1sp .

76:09:20·.38

165.2

•

Burn duration, sec .

Burn arc, deg

!::J.V, fps

DPS propellant used, Ib

Selenographic longitude of burn termination,deg E

Period of LPO, hr .

Altitude of pericynthion of LPO, n. mi.

Altitude of apocynthion of LPO, n. mi.

521

35.0

2398

16 421.3

159.9

3.02

58.3

707.0

Certain key parameters for this burn are plotted in figure 1}. RpCI'l.118e

the period of this orbit is approximately 3 hours, the SPS circulariza­tion burn was performed near the first pericynthion. If an additionalrevolution of tracking were required prior to circularization, thensome adjustments would have to be made in the postcircularization timeline, for example, reduce the rest period or change the time forinitiation of the DOl-day activities.

A brief summary of the SPS circularization. burn is presented inthe following table

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Time of ignition, hr:min:sec, g.e.t .

Selenographic longitude of burn initiation, deg W

Burn duration, sec

Burn arc, deg

7:10:24

176.6

43.1

2.8

Note that this burn would have the shortest SPS burn duration basedon the ground rules established for this alternate.

3.1.2 Alternate lb - LMtesting during translunar coast with CSM­only lunar orbital mission.- If the DPS LOI is not performed, then thesecond priority alternate is a CSM-only lunar orbital mission. The LMtesting would be performed during the translunar coast phase of themission. The following guidelines are recommended for alternate lb.

!J.V, fps .

SPS propellant used, lb .

10

620.9

3460••

l.to avoidwith thecritical

2.

The 1M testing should be performed early in translunar coastany trajectory perturbations caused by the activities associatedLM burns (separation , evasive ma.neuvers, etc.) close to thepre-LOI tracking.

No DPS burn will be performed.

3. An unmanned APS burn to depletion will be performed. This burnwill be targeted to place the ascent stage in a heliocentric orbit.

The e~rliest opportunity to perform LM checkout and testing is atapproximately TLI plus 19 hours. This time corresponds to the secondcre.r activity period after TLI. The nominal F mission LM checkout pro­cedures would be followed as closely as possible. Between approximatelyTLI plus 23 hours and TLI plus 25 hours, the DPS is staged and an unmannedAPS burn to depletion is performed. The APS burn targeting that isdesigned to place the ascent stage into a heliocentric orbit is veryinsensitive to the local horizontal burn attitude (fig. 12). It isarbitrarily recommended that the APS burn to depletion be made pitchedup 45° from the local horizontal. This attitude will provide good LMhigh-gain communications during the burn, and the ascent stage shouldeasily go into an orbit about the sun.

After the APS burn, the CSM returns to the nominal TLC time line.The CSM then will fly a CSM-only lunar orbital mission as described indetail in alternate 2.

3.1.3 Alternate lc - CSM/LM flyby.- If neither alternate la nor Ibcan be flown, then a CSM/LM flyby is flown with LM testing performed nearpericynthion. At PC minus 5 hours after LM checkout, a DPS burn ismade to raise pericynthion and to establish a free return to a CLA. The!J.V of the burn will be up to 500 fps (ref. 3).

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Shortly after pericynthion, the DPS is staged and an unmanned APSburn to depletion is performed. The APS burn is targeted along thevelocity vector, and the resultant 6V of approximately 3866 fps shouldplace the ascent stage in an orbit about the sun.

The APS burn to depletion was made near pericynthion for severalreasons.

1. The 1M was powered up and checked out and ready for a burn.

2. After jettison of the ascent stag~, the CSM is capable of agreater 6V for the subsequent MeC and for any weather avoidance problems.

3. All activities associated with an APS burn to depletion (1Mjettison, evasive maneuvers~ etc.) would be performed early in TEC andwould not be a source of trajectory perturbations. If these argumentsare not felt to be valid, then the option exists to keep the LM as acommunications backup until shortly peior to entry.

After sufficient tracking of the CSM, an SPS maneuver is made toreturn in the shortest possible flight time to a landing at the' PacificCLA.

3.2 Alternate 2: CSM-only Lunar Orbital Mission

If T, D, and E cannot be performed after TLI, then a CSM-onlylunar orbital mission will be flown. This alternate would also beused for a nonnominal TLI as described in alternate lb. The philosophywould be to follow the nominal Mission F time line work-reRt cycles.In lunar orbit, the times nominally scheduled for 1M checkout, rendezvous,and the APS burn to depletion would be used for landmark tracking andfor lunar surface photography. Included in the profile are severalrevolutions of MSFN tracking of the CSM in a low pericynthion orbit.Because alternate 2 is a radical departure from the nominal lunar orbit

,time line and because the lunar orbit events required :careful scheduling,a more detailed time line is being generated for it than were generatedfor the other alternate missions. This effort presently is beingcoordinated with FCSD.

The event time line is illustrated in figures 13 and 14. On theLOr daY, the major portion o~ the time available is allotted to lunarsurface photography. The photography consists of oblique target ofopportunity coverage to the north and to the south of the lunar ground­track. Photography is performed on revolutions 2, 3, and 4 (fig. 13).Photography takes up considerably less than half of revolution 3because of LOI-2 and the associated activities. Landmark tracking isperformed near the sunset terminator on revolutions 3, 4, and 5.

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Consistent with the nominal time line, a rest period is begun duringthe fifth revolution after LOr.

At a g.e.t. of approximately 94 hours or near the end of revolu­tion 9, the crew ends an 8-hour rest period. Two hours are allottedfor eating and for landmark tracking preparations. Near the beginningof revolution 11, landmark tracking is performed at CPl near the sunriseterminator, at CP2 near the subsolar point of the sun, and at Bl nearthe sunset terminator. These lan~~arks for the May 18 launch window are

ashown below.

Site designation Selenographic latitude Selenographic longitude

CPl 0053'N 170009'E

CP2 4°46'N 138°14'E

Bl 2°31'N 35°02'E

The sequence is repeated for revolutions 12, 13, and 14. Near the endof revolution 14, at 130.7° W, an SPS Dor maneuver is performed to in­sert the CSM into a 60- by 8-n. mi. orbit. The burn is targetedto place the pericynthion 15° up ra~ge of the landing site as in thenoreinal descent profile. Burn time is compared in figure 15 .,ith ~v

and resultant ~ericynthion altitude for the maneuver.

The CSM will remain in this orbit for three revolutions. On thefirst two pericynthion passes, landmark tracking is attempted on land­mark Bl. The geometry and attitude requirements associated with thesepasses over the site are shown in figure 16. Because the LOS rate tothe landmark is so rapid in this low orbit, it is not expected that thecrew will be able to perform more than two or three marks. The thirdpericynthion pass is used for vertical stereo photography.

Near the third apocynthion of the 60- by 8-n. mi. orbit, the CSMis circularized in 60-n. mi. orbit. A crew rest period which isstarted shortly thereafter consititutes a return to the nominal missionprofile.

aThe landmarks for subsequent days have not yet been selected. Thechoice is being coordinated between FCSD and the L~nar Mapping SciencesDivision.

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3.3 Alternate 3 - Contingency: LM NO-GO forUndocking and Rendezvous

3.3.1 Alternate 3a - DPS TEI.- If the LM checkout indicates tha.tthe LM was NO-GO for the undocking but that the DPS was GO for a burn,then basically the alternate mission would consist of a DPS TEl. Thedocked DPS capability compared with the CSM/LM weight is shown infigure 17. Normally, the CSM/LM weight after LOI-2 is approximately70 000 pounds. This weight will result in a DPS.6V capability of .approximately 2800 fps (fig. 9). Thetransearth flight time capabilitybased on this 6V is shown in figure 2 for the May launch windows.

The scheduling of the DPS TEl is a major consideration foralternate 3. A tradeoff exists betwee~ performance of a DPS TEl nearthe nominal time of the first DPS burn (DOl) and delay of the maneuveruntil after the landmark tracking scheduled for the third crew activityperiod in lunar orbit. This tradeoff, in turn, is a function of thebuildup of the supercritical helium pressure.

If the supercritical helium pressure were within the requiredlimits, then the time which was formerly used for rendezvous and forthe APS burn to depletion may be used for four revolutions of landmarktracking and for a DPS TEl. Based on the nominal flight plan, undocking

h m h m .occurs at 98 30 g.e.t. The APS burn occurs at 109 00 g.e.t., whlchallows approximately 10 hours and 30 minutes or a little over fivecomplete revolutions to perform landmark tracking and to perform TEl.An illustration of this profile is shown in figure 17.

After the 'four revolutions of landmark tracking, a real-timedecision would have to be made as to whether to stage the DPS and ~~­

main in orbit for next day's activities or to burn TEl. If the super­critical helium pressure is such that the previously outlined profilecannot be flown, it is recommended that the DPS be staged. The landmarktracking would be performed during the period outlined above and also.if required on the next day as in the nominal flight plan. As APSburn to depletion would be performed as in the nominal time line andlandmark tracking would be performed during the remainder of the crewday. The mission would then revert to the nominal timeline.

If the DPS were used for TEl, an unmanned APS burn to depletionwould he performed as soon as possible after TEl. This timing wouldallow the maximum amount of unperturbed tracking for the firGt trans­earth MCC which occurs at approximately TEl plus 15 hours. The APSis targeted along the velocity vector to maximize the chances of theascent stage going into an orbit about the sun. The MCC at TLI plus15 hours would be an SPS maneuver to return to earth in a minimumtransit time consistent with landing at 165 0 Wwithin entry velocitylimits.

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3.3.2 Alternate 3b - 1M APS burn to depletion in lunar orbit.-If the 1M were NO-GO for undocking and if the DPS were NO-GO for a burn,then an APS burn to depletion would be performed. The nominal time lineprocedures would be simulated as closely as possible until the nominaltime of the phasing burn. After this time, two additional revolutionsof 1M testing could be performed .. The DPS stage would then be jettisonedand the APS burn to depletion performed as in the nominal time line.

Based on the time used to perform further LM checkout beyond thenominal time line, landmark tracking could be performed in the timeperiod formerly used for rendezvous.

3.4 Alternate 4 - SPS TEl with Docked APS Stage

If there is a CSM communications failure in lunar orbit, thealternate would be to perform TEl and to keep the 1M as a communicationssystem. If the DPS is available, the DPS TEl could be performed asdescribed in alternate 3. If the DPS has been staged, then an SPS TElwith the ascent stage attached will be performed. Based on a nominalfuel loading after LOI-2, the SPS should have the capability to achieveapproximately 2800 fps, which would provide a slow transearth flighttime capability (fig. 3) .

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10 X 103987654321o

1

6

5

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f-'

~ 4

'01-~ . 1!1I11111111l11111111_lIll11IIII1IIIIIIIIIIIIIII1HlIIi!lIllll1ll1l111111111111111111111111111111111111~ 3 tifttllfliUttlMltltttlMl! II i1IIIItili It.11111 iffmrt1l3fOO1tmnm .a..VI

2

CSM-only 6.V remaining, fps

Figure 1.- CSM-only 6.V remaining tollowing a CSMjLM SPS burn.

•,~-:..

128

i?1f~~

124

;c "1+1 . :,"-;t

MAY IS) fifi ttlfit ~ ~.

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if'! ii~

120

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44 48 52 56

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><J

UJI-•

•

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Figure 2.- Mean TEI!::.V requirements versus TE flight time for May 1969 launch window .

• •

1400

1200+1";'

.lliffImr

.,

• •I =314.6 secsp

Thrust = 20 500 Ib

Flow ..Ie =65.162'~'~"!i

•

I

1000

'".e-~

><l

800

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400

Full CSM/LI\II

Approximate massin LPO

.<.,o-tjj-j

Approximate massafter TEl

I em....-.:I

200110 100 90 80 70 60

Mass, Ib

50 40 30 20 10 X 103

Figure 3.- 6.V resulting from a 40-second SPS burn versus mass at burn initiation.

, Fuel sumptank

Fuelstoragetank

.:.;.:::::;.;.;:::;:

Heliumtank

f['

Oxidizerstoragetank

Oxidizersumptank

~ :::::::=;=:::::.·.·.1

Heliumvalves

:::::;:;:;.:.:.:::;:'1

.:.:.1

~())

To SPS engineTo SPS engine

Figure 4 •- Flow of service propulsion propellants.

• • • • •

• • • • •

1000900800700400 500 600DPS!::.V,fps,

300200100o

200

1000_800

..c

.."'0

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a.. _\Q-400

Figure 5.- DPS!::.V capability based on low fuel loadings.

::r.:::h

320 x 103

280

240

~:~I~~~1~l;f~:1;:-:l-;:0·~1@Bill~~Eill;llllllillll~ll~illllill!llml~lll~~~~ ~ .. -._-

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32

28 I:

I

26 ~

~4.

24 FE

o 40 80 120 160 200 240 280 320 360 400

f\)o.....,-,

Ground elapsed time, sec

Figure 6.- TLI burn profile for May 17, 1969, 72° launch azimuth first injection opportunity •

• • • • •

• • • • •11 x 103

I10

9

8

7

III6.e-

uu::E

~ 5

4

3

2

1

030 40 50 60 70 80v 90 100 110 120 130 140 150 160 170 UK) x 103

i\)....

Apogee altitude, n. mi.

Figure 7. - Mee IN requi rements for various apogee ellipses resu Iti ng from premature 5-1 VB shutdowns.

I

22

•

•

•

•II

',!;lfll>H'Iiff ,;; ;:'1 ;tf 1,

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36

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28

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" "H T 1r ,

32

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54

58

82

66

62

94

70

78

86

90

Figure 8.- Propellant used and resultant mass following a dockedSPS burn.

•

• • • • •

52 X 10356606468727680848892

4.0

1.696

3.6

2.0

~ 3.2......><1V>

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Fi~re 9.- Docked DPS h.V capability for various masses at burn initiation.

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Figure 10. - Nominal SPS LOI for 72° launch azimuth first opportunity for May 17 launch window .

• • • • •

• • ~

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0

Burn time, sec

Figure 11 • - DPS LOI into a 6O-by 707-n. mi. orbit.

SC velocity, V = 5001 fpsParabolic velocity, V = 6292 fps

pParabolic deficit, V -V =1291 fps

p ---40---Altitude, h =113 565 n. mi.

Figure 12,- TLI + 25 hours APS burn to depletion,

••

•

••

• • • •fNf.......

•Landmark tracking

Target of opportunity photography

•

LOI-1~ 76 hr.LOI-2 -::: 80 .~5 hrBegin rest % B6 hr

/,un

I\)

-4

Figure 13.- Events sequence for LOI day.

....... Landmark tracking

H++H Vertical stereo photography

Nominal timeline events

Event Time, g.e.t.

1 - Awake and eat 94:00

Sun 2 - Begin LM checkout 95:00----- 3 - 001 99:54

4 - Phasing 101:06

5 - Insertion 103:03

6 - CSI 103:54. ~

f\)(J:) .

7 - CDH 104:52

8 - TPI 105:59

9 - Docking 106:40

10 - APS burn to depletion 109:04

• •

Figun 14. Event sequence for 001 day.

• • •

• • " • • •

76

Perilunealtitude

54321o

20

100

120

0-

10

50

60

..-E

40 t- 80.c..

a....s::::..

30 ~II)

Q) a.--0 60~ .......~.....

I\)~ \0Q)c:::s- 20 I- 40~Q)

a.

Burn time, sec

Figure 15.- SPS 001.

........ --- .... _-35°

LMKwo

(a) Perilune 16° behind landmark.

Figure 16. - Landmark geometry for 8-by 60-n. mi. lunar orbit.

• • ~ • • •

• •90

..

·i

•

--~-~~+++rt:

++

~

l:::t,~

m.1

p

-.~;.~r~.

rth

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•

;-.

'+"m

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'j.'-I-lii:W::; .++Ham +~

o 1 2Time, min:sec

3 4 5

(b) Elevation versus time for perilune of 60-by 8-n. mi. orbit.

Figure l6.-Continued.

\\

\\

\\\

\\\

\\

x

Localhorizontal

LMK

Cd Inertial altitude for landmark tracking a perilune of 60-by 8-n. mi. orbit.

Figure 16.- Concluded.

WNr-

• • .. • .. • • •

• • ..

•,. • • •

......... Landmark tracking

~un

TEl preparation

Figure 17 .- DPS TEl on 001 day.

Nominal timeline events

Event Time, g. e. t.

1 - Awake and eat 94:00

2 - Begin LM checkout 95:00

3 - 001 99:54

4 - Phasing 101:06\.}..J

5 - Insertion 103:03 '- .,?t:!.

6 - CSI 103:54 < .....7 - CDH 104:52

8 - TPI 105:59

9 - Docking 106:40

10 - APS burn to depletion 109:04

T,D,ANDE

EARTH ORBIT~"---~ ALTERNATE

CSM-ONLYLUNAR ORBIT

MCC FORCIRCUMLUNARMISSION

••

•NO

YES

SPS LOI CONTINUE a.-.-----.,NOMINAL MISSIONPROFILE

*WITHIN DEFINED GUIDELINES

DPS LOI-1SPS LOI-2

Ca) TLI contingency.

CSM/LMFLYBY

•

•Flow chart 1.- Decision logic for lunar alternate missions.

•

•

•

•

LUNAR ORBITLM C/O

35

NO STAGE DPS UNMANNEDAPS BURN TO DEPLETION

CONTINUENOMINAL MISSION

RETURN TONOMINAL TIMELINE

••

(b) Lunar orbit.

Flow chart 1.- Concluded •

36

REFERENCES

1. CSM/LM Spacecraft Operational Data Book, Volume I, Part 1 - Constraintsand Performance. SNA-8-D-027, May 1, 1968.

2. Lunar Mission Design Section: Preliminary Apollo Mission F TrajectoryData for the May 1969 Launch Window. MSC IN 69-FM-27, February 4,1969.

3. Zeiler, Kenneth T.: Apollo Mission F ~v Requirements for a Flybyto Hawaii at LOI - 5 Hrs. MSC memo 69-FM52-46, February 24, 1969.

••

•

•

•

••