Lunar Module Descent and Ascent

8/8/2019 Lunar Module Descent and Ascent

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N A S A T E C H N I C A L N O T E

APOLLO EXPERIENCE REPORT -MISSION PLANNING FOR LUNAR MODULEDESCENT AND ASCENT

1 by Floyd V, BennettMunned Sucecrdft CenterHozlston, Texus 77058

N A T I O N A L A E R ON A U TI C S A N D S PA CE A D M I N I S T R A T I O N W A S H I N G T O N , D. C. J U N E 197


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I. Rewrt No. 2. Government Accession NoNASA TN D-68464. Title and Subtitle

A P O L L O E X P E R I E N C E R E P O R TMISSION PLANNING FOR LUNAR MODULE

3. Recipient,'s Catalog No.

6 Reoort DateJune 19726. Performing Organization Code

Floyd V . Bennett , MSCDESCENT AND ASCENT

7. Author(sJMSC S-295

8. Performing Organization Report No.

9. Performing Organization Name and AddressM a nne d S pa c e c r a f t C e n te rHouston, Texas 77058

2. Sponsoring Agency Name and Address

10. Work Unit No.076-00-00-00-72

11. Contract or Grant No.

13. Type of Report and Period Covered

Na t ional Ae r o na u t i c s a nd S pa c e Adm in i s t r a tionWa sh ing ton , D. C. 20546

f this page)

14. Sponsoring Agency Code

21. 'NO. of Pages 22. Price*48 $3.00

I5. Supplementary Notes: h i s Apo l lo Exp e r i e nc e R e p or t , because, in h i s j udgm e n t , use of SI Uni t s wou ld im pa i r the usefulnes,3f the report or r e s u l t in excessive c o s t .

T h e MS C D i r e c t o r w a i ve d t h e u s e of the I n t e r na tiona l S ys t e m o f Un i t s (SI) o r6. Abstract

The p r e m is s io n p la nn ing , t he r e a l - t im e s i tua t ion , a nd the post f li gh t a na lys i s f o r t he Apo l lo 11l u n a r d e s c e n t a n d a s c e n t ar e d e s c r i b e d i n t h i s r e p o r t . A c o m p a r i s o n b e t w ee n p r e m i s s i o n p l a n -n in g a n d a c t u a l r e s u l t s is i nc lude d , A na v igat ion c o r r e c t i on c a pa b i li t y , de ve lope d f r om Apo llo 11pos t fl i gh t a na lys i s , wa s u se d suc c e s s f u l ly on Apo llo 12 o p r ov ide the f i r s t p inpo in t l a nd ing . Ane xpe r i e nc e sum m a r y , wh ic h i l l u s t r a t e s t yp ic a l p r ob le m s e nc oun te r e d by the m is s ion p l a nne r s ,is a l s o inc lude d.

19. Security Classif. (of this report)None

20. Security Classif. (None

18. Distribution Statement

For sale by he National Technical Information Service, Springfield, Virginia 22151


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I

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CONTENTS

Page112

SUMMARYINTRODUCTION. . . . . . . . . . . . . . . . . . . . . .PREMISSION PLANNING . . . . . . . . . . . . . . . . .I . . . . . . . . , .

Descent Planning . . . . . . . . . . . . . . . .Ascent Planning

* 4 ' * ' . * * *~ I~

I I

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .REAL-TIMEANALYSIS. . . . . . . . . . . . . . . . . . ( . . . . . . . . . . .

Descent Orbit Insertion . . . . . . . . . . . . . . . . . . . . . . . . . . .Powered Descent . I . . . . . .II. . . . . . . . . . . . . . . . . . . . . . . . ' I

IAscent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .POSTFLIGHT ANALYSIS . . . . . . . . . . . . . . . . . . . . . I . . . . . . .I

I/Apollo 1 1 Descent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Apollo 11 Ascent . . . . . . . . . . . . . . . . . . . . . . . .; . . . . . . .1 IAPOLLO 12 MISSION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Apollo 12 Planning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Apollo 12 Posff light Analysis I. . . . . . . . . . . . . . . . . . . . . . . . .

MISSION-PLANNING EXPERIENCE . . . . . . . . . . . . . . . . . . . . . . .System Design Specifications . . . . . . . . . . . . . . . . . . . . . . . . .Definition of System Pe rf ormanc es . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ystem InterfacesMission- Planning Flexibility . . . . . . . . . . . . . . . . . . . . . . . . . .Experience Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONCLUDING REMARKS

313161616202 121252626293132333738394 04 2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .EFI~RENCES

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TableI

I1

I11

IVV

Figure123

iI .1I

TABLES

APOLLO 11 PREMISSION POWERED-DESCENT EVENTSUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .APOLLO 11 PREMISSION DESCENT AV AND PROPELLANTREQUIREMENTS . . . . . . . . . . . . . . . . . . . . . . . . . . .APOLLO 11 PREMISSION A S C E N T A V AND PROPELLANTREQUIREMENTS . . . . . . . . . . . . . . . . . . . . . . . . . . .APOLLO 11LUNAR-DESCENT EVENT TIMESAPOLLO 11 ASCENT SUMMARY

. . . . . . . . . . . . .(a) Events . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .(b ) Inse rtion conditions . . . . . . . . . . . . . . . . . . . . . . . . . .(c) Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

FIGURES

Lunar module descent . . . . . . . . . . . . . . . . . . . . . . .bera t ion al phases of powered descent . . . . . . . . . . . . . . . . .Lunar module body axes and LR antenna axes

Ia) Lunar module body axes . . . . . . . . . . . . . . . . . . . . . . .c) Landing radar position 1 (used in braking phase) . . . . . . . . . .:d) Landing ra dar pos,ition 2 (used in approach and landingp h a s e s ) . ~ .. . . . . . . . . . . . . . . . . . . . . . . . . . . .Lunar module forward window . . . . . . . . . . . . . . . . . . . . . .Basic LM descen t guidance logic

I

b) Landing ra da r antenna axes . . . . . . . . . . . . . . . . . . . . .I

. . . . . . . . . . . . . . . . . . . .Target sequence fo r abtom atic-descent guidance . . . . . . . . . . . .Pre miss ion Apollo 11 LM powered descent . . . . . . . . . . . . . . .

iv

Page

9

12

1522

252526

Page33

55556678


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L , _ f i "I!I

8 Premission Apollo 11 time history of thrus t and attitude(a) Thrust . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.; . .7

I

(b) Attitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .I . . . . . . . I * *I. . . . . . . . . . . . . . . . . . . . . . . . . .anding phase , . .;.. . . . .redi cted Apollo 11 landing dispersions 7 ' ' * * ' * *1011 1111

1213

I i. . . . . . . . . . . . . . . . .remission Apollo 11 LM ascent ./ .;.~Pr emi ssi on Apollo 11 vertical-r ise phase . . . . / . . . . . . . . . . .1

14

1314

1516

1414

Pr emi ssi on Apollo 1 1 orbit-insertion phasePredicted Apollo 11 asc

. . . I . . . . . . . . . . .. . . . . I I. . . . . . . . . . .1 1Effect of position e r r o r on velocity comparison . .I. . .~. . . . . . .

17 Guidance thrust command as a function of horizonjal velbcity . . . . . . i18 Landing radar altitude updates . . . . . . . . . . . . . . . . . . . . . ~ 19I I19 Altitude as a function of altitude rate during powered descent . . . . . j 20

20 Apollo 11 landing site * I . * * * ' * " 2021 Apollo 11 approach phase . . . . . . . . . . . . . . . I . . . . . . . 2322 Apollo 11 landing phase . . . . . . . . . . . . . . . . . . . . . . . . . 23

I. . . . . . . . . . . . . . . . .

I . . . . . . . . . . . . . . .3 Apollo 11 groundtrack fo r the landing phase 24. . . . . . . . . . . . .4 Apollo 11 attitude profi le fo r the landing phase 2425 Apollo 11 altitude as a function of altitude rate for the landing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .hase 2426 Apollo 11 landing-phase events. 24. . . . . . . . . . . . . . . . . . . .27

28

Landing site update capability during braking phase(a) Throttle margin time 27(b) Change in chara cte ris ti c velocity . . . . . . . . . . . . . . . . . . 27. . . . . . . . . . . . . . . . . . . . . . . .Predict ed Apollo 12 landing dispersions . . . . . . . . . . . . . . . . . 28

V


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Figure2930

3132

33

3435

Variation of AV with landing-phase velociti es . . . . . . . . . . . . .The AV require ments fo r down-range redesignations at a 4000-footaltitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Comparison of Apollo 11 and Apollo 12 LPD profiles . . . . . . . . . .Apollo 12 window views 30 seconds a ft er high gate (al titude, 4000 feet)(a) Right-hanh window view taken with onboard 16-millimeter(b) Lunar module altitude above the landing site as a function of

IApollo 1 2 approach phase(a) Landing point designator angle as a function of sur fac e distance

to the landing site . . . . . . . . . . . . . . . . . . . . . . . . .(b) Computer recons truction of commander' s view . . . . . . . . . . .

1ca mer a (camera tilted 4 1' o the horizon) . . . . . . . . . . . .surface distance to the landing site . . . . . . . . . . . . . . . .

I Apollo 12 groundtrack for the landing phase . . . . . . . . . . . . . . .1

Attitude as a functioniof up-range distance for the Apollo 12 approachand landing phase s 1. . . . . . . . . . . . . . . . . . . . . . . . . . . .

Page28

2828

2929

303030

31

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' i

f

II

APOL L O EXPER IEN C E R EPOR T 1

! I II~

\ S U M M A R Y II \ , > 1

Premi ssi on planning, real -time analysi s of miss idn events, and posffliglys is are desc ribed for the lunar module descent and asce nt phases of the Apollmission, the first manned lunar landing, and for the Apollo 1 2 mission, the fircapability was provided for the Apollo 1 2 descent. Flight res ult s for both misishown to be in agr eement with premi ssion planning. A summary of mission-p!experience, which il lu st ra te s typical prob lems encountered by ithe mission plaiis als o included in this, repo rt.

I '!' \

point lunar landing. Based on the Apollo 11 postflight analysis , a navigation c(anal-11

t\ pin-rectionons arenningers ,I

II N T R O D U C T I O N II

Premi ss ion planning fo r Apollo lunar module (LM) descent and ascent sta rted in1962 with the decision to us e the lunar orbit rendezvous (LOR) technique for the Apollolunar-landing mission ( ref. 1)J The LOR concept advanced by Houbolt and others isdefined in refe renc es 1 and 2. The technique allowed optimization of both the des ign ofLM yste ms and trajectories f o r orbital descent to and ascent from the lunar surf ace.',The LM descent wa s designed to be accomplished in two powered-flight maneu-vers : the descent orbit insertion (DOI) maneuver and the powered-descent maneuver.The DO1 maneuver, a short o r iiulse-type tra nsfe r maneuver, is performed to re -duce the orbit alt itude of the LM from the command and service module (CSM) parkingorbit to a lower altitude fo r efficidncy in initiating the longer, mor e complex powered-descent maneuver. The basic trajectory design for the powered descent wa s dividedinto thre e operational phases : an init ial fuel-optimum phase, a landing-approach tran-sition phase, and a final translation and touchdown phase. The initial trajecto ry anal-ys is which led to this design was performed by Bennett and Price (ref. 3 ) . Inreference 4, Cheatham and Bennett provided a detailed descrip tion of the LM descent

*The mat eri al presented in this repor t, with the exception of the sect ion entitled"Mission-Planning Experience, " was previously published in NASA TM X-58040.


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design strateg y. This description illustr ates the complex interactions among sy st em s(guidance, navigation, and con tro l; propulsion; and landing radar), crew, trajectory,and operational cons train ts. A more de tailed descr iption of the guidance, navigation,and control system is given by S ear s (re f. 5). As LM sy ste ms changed f ro m designconcept to hardware, and as operational constraints were modified, it became neces-sa ry to modify o r reshape the LM descent traject ory; however, the basic three-phasedesign philosophy was retained.The LM ascent was designed as a single powered-flight maneuver to ret ur n thecrew from the lunar surface, o r fro m an aborted descent, to a satisfactory orbit fro mwhich rendezvous with the CSM could be performed . The basic trajec tor y design forthe powered ascen t was divided into two operational phases: a vertical- rise phase forsurface clearance and a fuel-optimum phase fo r or bi t insertion. Thus, the ascentplanning was more straig htfo rward than the descent planning and, because of the lackof a lunar atmosphere , l e s s complex than earth-launch planning.The purpose of this report is to descr ibe the p remission operational planning f o rLM descent and ascen t; that is, to describe the bridge from design planning to flight-operation stat us. A discussion of the primary c'riteria which precipitated the plan fo rthe Apollo 11 mission, a comparison of the r eal-tim e mission events with this plan, a

discussion of the postflight ana lysi s of the Apollo 11 mission and its application to theApollo 12 and subsequent miss ions, and a brief postflight discussion of the Apollo 12mission are included in thi s re po rt . In addition, a section on mission-planning expe-rience is included to provide insight into typical prob lems encountered by the missionplann ers and the solutions that evolved in to the final operational plan.The author wishes to acknowledge the ass ista nce of the mem ber s of the LunarLanding Section of the Landing Analysis Branch (Mission Planning and Analysis Divi-sion), particularly, W. M . Bolt, J. H . Alphin, J. D. Payne, and J. V . W e s t , whocontributed to the generation of the data presented in this report.

with Ecausebeen 1and pl

neuvenent Echaraprovic

2

' P REM IS S ION PLANN NGI I

-emission planning entails the integration of mission r equir ements o r objectivestem and crew capabilities and constraints. The integration is time varying be-?ither mission req uir ehe nts nor system performances remain static. This has-titularly true of the d M descent and ascen t maneuvers, which wer e in designning fo r 7 years . II I

thi s section, the final evolution of the planning for the descent and ascent ma-for the Apollo 1'1mission will be descr ibed. A brief descr iption of the per ti-tems, the guidance logic, the operational-design phases, the trajectoryIristics, and the AV and propellant requir ement s for each maneuver isx.~

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Descen t P lann ingThe LM descent from the CSM park-ing orbit (approximately 62 by 58 nautical

miles) is illustrated in figure 1. After theLM and the CSM have undocked and sepa-rated to a safe distance of se ve ra l hundredfeet, the LM performs the DOI, which isthe first and simplest of the two descentmaneuvers . The DOI, which is a shortret rog rad e maneuver of approximately7 5 fps, is performed with the LM descentengine and is made at a positbi t 180' fr om powered desce(PDI), which is the second de II

PhaseBraking

ApproachLanding

Figure 1.1- Lunar module descerer . The purpose of the DO1 is to reduceefficientlv (with Hohmann-type transfer) I

Design criteriaHigh gate Crew visibilityLow gate Man ual control

Initial eventPDI Minimi ze propellant usage

the orbit "altitude fr om approximately60 nautical mil es to 50 000 feet in prepara tion for PDI. Performance of continuopowered descent fro m altitudes much gre ater than 50 000 f k e t is inefficient, and :at lower than 50 000 feet is a safety hazard (ref, 3). The DO1 isldescribed i n theonerational trai ectorv documentation at the NASA Manned Spaceckaft Center and !

I

tby

)

.kingI n

,.SPDI3d:lscussed furt


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System descriptions. - The success of the LM powered descent depends on thesmooth interaction of s eve ral sys tem s. The pertinent systems are the primar y guid-ance, navigation, and control sys tem (PGNCS); the descent propulsion system (DPS);the reac tion con tro l syste m (RCS); the landing rada r (LR); and the landing point desig-nator (LPD). A detailed description of each system and its performance characteris -tics is given in reference 6. A brief descrip tion of each sys tem follows.

tions,phase

The PGNCS consist s of two major sub syst ems: an inertial meas urement unit(IMU) and a computer. The IM U is the navigation senso r, which incorpo rates acce ler-om eter s and gyr os to sens e changes in LM velocity and attitude. The IM U sends thisinformation to the computer, which contains preprogramed logic fo r navigation, fo rcalculation of guidance commands, fo r sending steer ing commands (by means of thedigital autopilot (DAP)) to the DPS and the RCS, fo r processing LR measurements ofLM range and velocity re lative t o the lunar s urface, and fo r display of information tothe crew . The crew con trol s the mode of computer operation through a display andkeyboard (DSKY) assembly. A description of the guidance logic is given in a subsequentsection, and a complete description of the guidance, navigation, and control logic canbe found in reference 7.

as shown in fig ures 3(c) and 3(d). Position 1 (fig. 3(c)) is used in the brakingof the descent when the LM is orient ed nea r the horizontal. Position 2 (fig. 3(d))

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Location of LR

I Bores'ght' (b) Landing ra da r antenna axIi/I

I

'B 'AI

II

z A(c ) Landing radar position 1 (used inbraking phase).

5 .

(d) Landing rada r position 2 (used i napproach and landing phases ).Figure 3 . - Lunar module body axes and LR antenna axes.

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The final system to be described is agrid on the commander's forward windowcalled the LPD (fig. 4). The window ismarked on the inner and outer panes toform an aiming device or eye position.During the approach and landing phases,the computer calculates the look angle (rel-ative to the forw ard body axis Z B) to thelanding site and displays it on the DSKY.The commander can then sight along theangle on the L PD (ze ro being along bodyaxis Z ) to view the landing area to whichhe is being guided. If the commander de-s i r e s to change the landing area, he canmake incremental changes inplane o r c rossrange by moving the hand con trolle r in theappropriate direction to provide input to t h ecomputer. C r o s s - range position is changedin 2 increment s, and inplane position ischanged in 0. 5" increments. A detaileddescripti on of the guidance logic is given i nreferences 7 and 8.

B

Guidance logic. - ,Thebasic L M de-scent guidance logic i q defined by an ac-celeration command which is a quadraticfunctidn of t im e and is, therefore, termedquadratic guidance. A simplified flow charof quadratic guidance is given in figu re 5.The curren t L M position and velocity vec-to r s 6 and I? a r e determined fro m thenavigation routine. The des ired (or target)positidn vector I$ velocity vector tsD,acceleiation vector A and ,down-rangecomponent of je rk jthe std red memory. (Jerk is the time

IAD'I are obtained fromDZ

(looking outboardl

Figure 4. - Lunar module forwar dwindow.

Throttle DA P

Figure 5. - Basic LM descent guidancelogic.

d e r i v a p e of acceieration. ) :The down- range (horizontal) components of t he se statevecto rs (current and desir ed)iar e used in the jerk equation to determine the time to go(TGO);! that is , the time to go from the curren t to the de sir ed conditions. If th e TGO,the current state vector, and 'the desir ed s tat e vector are known, then the commandedacceleration vector A is determined from the quadratic guidance law. Note that theacceler ation- command equation yields infinite commands when the TGO rea ch es zero .For this reason, the targeting is biased such that the de sir ed conditions a r e achieved

A

CIIreaching zero. By using spacecraft mass M, calculating the v ectorce between the commanded acceleration and the acceleration of lunar gravity A ,

Newton's law, a commanded thrus t vector TC can be found. The6


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IImagnitude of TC is used to provide automatic throttling of the DPS. When:the tbrottlecommands exceed the throttl e region of the DPS (10 to 60 percen t of design thrust ),maximum thrust (FTP) is applied. The vecto r direction is,use d by the DAP to orie ntthe DPS thrus t by either tr im gimbal attitude commands o r RCS commands to re oh en tthe entire spacecraft.

I ,

I

During the powered descent, the guidance computer provides se ve ral se qu eh ia lpro gra ms (P63 to P67) fo r guidance and control operations. A description of each pro-gra m follows. A complete de scrip tion of the desc ent guidance logic and guidancelmodesis given in references 7 to 9. The first program is P63, entitled "Braking Phase Guid-ance. '( Pr ogr am 63 contains an ignition algorithm and the basic;guidance logic. Theignition logic, which dete rmines the time for the c rew to ignite the DPS fo r PDI, fsbased on a stored, preselected surface range to th e landing site.$ After descent-engineignition, the bas ic guidance logic is d to steer the L M to the +s ire d conditions forthe beginning of the approach phase. stat ed previously; the targets are selecdedwith a bias such that the des ire d conditions are achieved pri or to the TGO reaching zero.When the TGO reac hes a prese lecte d value, the guidance pro gra m switches automati-cally fr om P63 to P64, which is entitled ('Approach P hase ' Guidance. " Pr og ra m 164contains the sa me ba sic guidance logic as P63, but a new set of targets is selected toprovide traj ect ory shaping throughout the approach and landing phase s and to estqblishconditions f or initiating an automatic verti cal descent fr om a low altitude to landing. Inaddition, P64 provides window-pointing logic for the LPD operaqon. That is, the land-ing point will be maintained along the LPD grid an the comma nder' s window. Dutin gthis time, the crew can make manual inputs with the attituae hand con trolle r to changeincrementally (down range or cross range) the intended landing site and remai n in auto-,mati c guidance. (See the sect ion entitled "System Descriptions./") ~

When the TGO reac hes a preselectedvalue, the guidance progra m switches auto-matically from P64 to P65, which isentitled "Velocity Nulling Guidance."Program 65, which nulls all components ofvelocity to prese lecte d values, is used foran automatic vertic al descent to the sur -face, i f des ire d. No position control isused during this guidance mode. The se -quencing fo r automatic guidance is illus-

~

20=4

trated in figure 6. Range'Guidance P6 5 IS velocity nulling only

Pr og ram 66, entitled "Rate of De- ( 1 . e . . no position target1scent, ( ' and program 67, entitled "ManualGuidance, " a r e optional modes which canbe used at crew d iscre tion (manually calledup through the DSKY) at any time duringthe automat ic guidance modes (P63, P64,or P65). During P66 operation, the crew control space craft attitude, and the computercommands the DPS thro ttle to maintain the des ired altitude rate. This ra te can be ad-justed by manual inputs from the crew and is normally entered late in P64 operation(near low gate) pr io r to P65 switching fo r manual control of the final touchdown position.Pr og ra m 67 maintains navigation and display operations fo r complete manual control ofthe throt tle and altitude. Normally, this mode is not used unless P66 is inoperative.

Figure 6. - Target sequence forautomatic-descent guidance.

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formesurfacdatingis exp

Braking phase. - A scale drawing of the LM powered descent for the Apollo 11mission is given in figure 7. The intended landing area, designated Apollo site 2, inthe Sea of Tranquility is centered at latitude 0.6" N and longitude 23.5" E. The majorevents occur ring during the braking phase ( illustrated in figure 7 and tabulated intable I) are discussed as follows. The braking phase is initiated at a preselected rangeapproximately 260 nautical miles from the landing site near the perilune of the descenttr an sf er orb it (altitude of approximately 50 000 feet). This point is PDI, which coin-cid es with DPS ignition. Ignition is preceded by a 7.5-second RCS ullage burn to settlethe DPS propellants. The DPS is ignited at tr im (10 percent) throttle. This throttlesetting is held for 26 seconds to allow the DPS engine gimbal to be alined (or trimmed )through the sp acecraft center of gravity before throttling up to the maximum or fixedthrottle position. The braking phase is designed for efficient reduction of orb it velocity(approximately 5560 fp s) and, therefore, uses maximum thru st for most of the phase;however, the DPS is throttled during the final 2 minute s of th is phase fo r guidance con-tr ol of dispersion s in thrust and trajecto ry. As stated earlier, the DPS is throttleableonly between 10 and 60 percent; therefore, during FTP operation, the guidance is tar-geted such that the commanded quadratic accelerat ion, and consequently the commandthrust, is a decreas ing function. When the command dec rea ses to 57 percent, a3-percent low bias, the DPS is throttled as commanded (illustrated by the time historyof commanded and actual thrus t shpwn in fig. 8(a)). The thr ust attitude (pitch) profileis shown in figure 8(b). Ea riy in the descen t, orientation about the thr us t axis is bypilot discretion. The Apollo 11 crew oriented in a windows-down pttitude fo r visualground tracking as a gr os s navigation check. Rotation to a windows-up attitude is per-

I at an altitude of approximately 45 000 feet, so that the LR can acquire the lunar3 to update the guidance computer es tima te s of altitude and velocity. Altitude up-is expected to begin a t' an altitude of approximately 39 000 feet; velocity updatingcted to begin at approximately 22 000 feet.

0@-io'-

I I I11.54 8 7Th

8

618 sec

4 :Begin

I East lunarlongitude

Altitude12-n. mi.

28"6 5

circled letters correspond to the events listed in table 1III

I -1IFigure1 7. Premission Apollo 11 LM powered descent.

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TABLE I. - APOLLO 11 PREMISSION POWERED-DESCENT E V E N T SUMMI

Event

UllagePowered descent initiationThrottle to maximum thr ustRotate to windows-up positionLR altitude updateThrottle recoveryLR velocity updateHigh gateLow gateTouchdown (probe contact)

TFI,min: sec(a)-0: 7

0: 000: 262: 564: 186: 246: 428: 26

10: 0611:54

"Time fr om ignition of the DPS.

Inertialvelocity,fPS

556055294000306514561315,506b55(68)

b-15(0)

Altituderate,fPS

-4-3

- 50I-89

-1061

I-127-145-16

I

- 5

Altitude,f t

48 81448 72544 93439 20124 63922 644

7 515 ~51212

I

IbHorizontal velocity relative to the lunar surface. I,

M lo3I -Actual Commanded---16 r\. 30 dispersionI throttle recovery

e ..1' -'\,H\ \

r \'Ti4

031

1572253642 394399537561766775

-m2>Ee-Sco.

Time from ignition, minTime f r o m iqnition. min

(a) Thrust. (b) Attitude.Figure 8. - Premission Apollo 11 tim e history of thru st and attitude.

9


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upon making surface contact, activate alight which signals the c rew t o shut downthe DPS manually, whether automatic ormanual guidance is being used . The landing-phase trajectory is shown under automaticguidance in figure 10.Premiss ion estimates of dispersionsin landing position a r e shown in figure 11.These dispersions, which are based on aMonte Car lo analy sis, include all knownsys tem performances as defined in re fer -ence 6. Based on thi s analy sis, the99-percent-probability landing ellip se wasdetermined to be f 3.6 nautical miles in-plane by & 1. 3 nautical miles c ro ss range.

Cross range,n. mi .

Forward window view

loo0800

e7 6005'B 400

200I

0 4w 800 1200 lboo 2000 2400I I I I I

Range, ft0 .1 . 2 . 3 .4

Range, n. mi .I 1 I I I11:54 11:OO 10:40 1o:zo 1o:oo

TFI . rnin:sec

Figure 10. - Landing phase.

. -

Figure 1.1.- Predicted Apollo 11 landing dispersions.The AV and propellant requirements. - The AV and propellant requireme ntsa r e determined by the nominal trajectory design, contingency re quirem ents, and dis-persions. Consequently, these requir eme nts have undergone continual change. Thefinal operation requirements are given in table 11. The required 6827-fps AV is e s -tabl ished by the automat ically guided nominal. In addition, 85 fp s is added to assure2 minutes of flying time in the landing phase, that is , below an altitude of 500 feet.The automatic guidance requi red only 104 seconds of flying time for the landing phase.Also, a 60-fps A V is added for LPD operation in the approach phase to avoid largecr at er s (1000 to 2000 fee t in diamet er) in the landing ar ea . Contingency propellantallotments a r e provided fo r failure of a DPS redundant propellant flow valve and forbia s on propellant low-level-light operation. The valve failure causes a shift in thepropellant mixture ratio and a lower thrust by approximately 160 pounds, but otherwise,DPS operation is sati sfac tory. The low-level light signifies approaching propellant de-pletion; therefore , a bias is used to protect against dispersi ons in the indicator. If thelow-level light should fail, the cre w use s the propellant gage reading of 2 percent r e -maining as the abort decision indicator. The light se ns or provides more accuracy and

i

\

11

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is ther efor e pre fer red over the gage reading. The ground flight con trol lers c all outtim e fro m low-level light ''on'' to inform the crew of impending propel lant depletion fora land-or-abort decision point at least 20 seconds before depletion. This proced ureallows the crew t o start arr es tin g the altitude rate with the DPS pri or to an abort s tageto prevent surfac e impact. The allowance for disp ersions is determined fro m the MonteCarl o a nalysi s mentioned previously. A s can be seen in table II, the AV and propel-lant requirements are satis fied by a positive margin of 301 pounds. Thi s marg in canbe converted to a n additional hover o r t ransla tion time of 32 seconds.

TABLE 11. - APOLLO 11 PREMISSION DESCENT AV ANDPROPELLANT REQUIREMENTS

Item

UnusAvaiNomDispPadContE1RIRM

M a r-

tbleLble for AV I

. Iial required for AV (6827 ps)rsions (-30) I

ngencie s Igine-valve malfunctionl

~dline low-level sensor~designation (60 ps ) I

nual hover (85 ps) Iin I

I

I

,I

Propellant required,lb

- -75.4

250.5- -

l a , ,16 960.9292.0- -

64.768.7102.9144.0- -

Propella nt remaining,lb

18 260.518 185.117 934.617 934.6

973.7681.7681.7

617.0548.3445.43.01.301.4

7051.2 pounds of fuel land 11 209.3 pounds of oxidizer.Fuel offload of 75.4 pounds to minimize malfunction penalty.

13

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A s c e n t PlanningA sketch of the LM ascelunar surfa ce is given in fig ure 12 . Theascent has a sing le objective, namely, t o

achieve a satisfactory orbit f ro m whichrendezvous with the orbiting CSM can sub-sequently be perfo rmed. Nominally, in-sertion into a 9- by 45-nautical-mile orbit ,at a tr ue anomaly of 18' and an altitude of60 000 feet, is des ired. The time of lift-off is chosen to provide the pro per phasingfo r rendezvous. A desc ription of thepowered ascent, not the choice of targ etingfor rendezvous, is the subject of thi ssection.Figure 1 2 . - Prernission ApollcI LMascent .

System descriptions. - 'Only thr ee pertinent sy ste ms ar e required fo r ascethe PGNCS and RCS, which have already been described, /and tly ascent propulssystem (APS). The APS, unlike the DPS, is not throttleTble and does not have 2gimbal drive , but provides a constant thru st of approximately 3500 pounds throuthe ascent (ref. 6). Engine throttling is not required durfng ascent, because dorange position control is not a target requirement; that is, only/ altitude, velocitorbit plane are required for targeting. ' This thrust can bb enhahced slightly (byimately 100 pounds) by the RCS attitude control. The ascent DAP logic is suchbody positive X-axis (along the thrus t d irection) jets a r e fired fo r attitude contrIascent. IA fourth sy ste m, the a bor t guidance sys tem (AGS), should also be mentioiThe AGS is a redundant guidance syst em to be used fo r guidance, navigation, ar

\\\ un

tI. orbit)

11

t -Intr imhoutn-, andtpprox-iat only1 during

!d.con-tro l for ascent o r ab or ts in the eve& of a failure of the PGNCS! . The AGS has its owncomputer and uses body-mounted se ns or s instead of the inertia l sen so rs as used in thePGNCS. A deta iled description of the AGS is given in references 11 and 1 2 .Operational phases. - The powered ascent is divided into two operational p hases:vertical ris e and orbit insertion. The vert ica l-ri se phase is required to achieve ter-rai n clearan ce. The trajec tory f or propellant optimization takes off along the lunarsurfa ce. A description of traject ory p ara me ter s and LM attitude during the vertical-ri se phase and during the transition to the orbit-inse rtion phase is shown in figure 1 3 .The guidance switches to the orbit-insertion phase when the radial rate becomes 40 fps.However, because of DAP ste er in g lags , the pitchover does not begin until a radial rateof approximately 50 fps is achieved. This delay means that the vertic al-ri se phase isterminated 1 0 seconde after lift-off. Also, during the verti cal ri se , the LM body

Z - a x i s is rotated to the de si re d azimuth, which is normally in the CSM orbit plane.The orbit-insertion phase is designed for efficient propellant usage to achieveorb it conditions for subsequent rendezvous. The orbit -insertion phase, the tota lascent-phase performance, nsertion orbit param eters , and onboard displays at inser-tion are shown in figure.14 The onboard-display values reflect the computer-estimated

values. If required, yaw steering is u s ring the orbit-insertion phase to maneuverthe LM into the CSM orb it plane or into a plane parallel with the CSM orbit. In the13


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nominal case, no yaw stee ring is require d. The nominal ascent burn time is 7 minutes18 seconds w i t h a 3a disper sion of ? 17 seconds. The trajecto ry dispersions areplotted in figu re 15. The ascent guidance logic is discus sed in the following section.

1(

14

"2--,o 12-.-".-

b- 10

8

64

O 2

Fif

guidarbecau,For tlsimp11stage,verticto 40descei

I Guldanceswitch

is a lit e r m i14

- .

- 20-640

- 60-480

-400 ral=

-320

- 240- 160

-8 0

100c b ,160 80 0 -80 -160Down-range powtion. ft II

Ire 13. - Prem iss ion Apollo 11vert ical-r ise bhase.II

hida nce logic. - iThe asc ent-:e logic commands onlv attitude.

Ascent burn out coast to45-11, mi. apolune

Orbit-insertion phase

End vertical rise

Total ascent:Burn time = 7 min 18 secAV required = 6060 fpspropellant required - 4934 Ib

Insertion orbit parameters Onboard displaysat insertionHeight at perilune, h_ - 55 OOO 11PHeight at apolune. ha - 45 n. mi.True anomaly, , 180Flight path angle, Y = 0. f V = 5535.6 fpsh li 60 085.4 ftti - 32.2 fps

Figure 14. - Prem iss ion Apollo 11orbit- inser ion phase.Insertion -lo3

6 0 -

50-c

$40--=30-

Terrainestimates, ,

Schmidt7 ---J30

0 20 40 60 80 lo o 120 140 160 180 200Range, n. m i.Figure. 15. - Predic ted Apollo 11ascent dispersions.e no engine throttling l's required.

? vertical-r ise phase, the logic is. The initial attitude is held for 2 seconds in orde r to clear the L M descentthe attitude is pikhed to the vertica l while rotating to the desire d azimuth, andI-ri se- phas e ter'mination occ urs when the altitude rate is grea ter than o r equalIS pward, or when the altitude is grea ter than 25 000 feet (used for aborts fr omt.

i

'he insertion-phase guidance logic is defined by an acce lera tion command whichear function of ti me .and is, therefore, termed linear guidance. The TGO ised as a function of velocity to be gained; that is, the difference between th


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I

vated to cut off the APS at the desired time. II I

Propellant required,lbtem

,1 II

TABLE ILI. - APOLLO 11PREMISSION ASCENT AV ANDPROPELLANT REQUIREMENTS

'1

Propellant remaining,lb

I UnusableAvailable for AVNominal required for AV ( 6 0 5 5 .7 fps)Dispersions (-30)PadContingencies

Engine-valve malfunctionPGNCS o A M switchover (40 fps)

56 . 3 5167 .4-- 5 1 6 7 .4

4 9 6 6 .7 200.766 . 7 1 3 4 . 0_ - 1 3 4 .01 8 . 8 1 1 5 . 22 3 .8 9 1 .4

1 I

Abort from touchdown(AW = + 1 1 2 .9 lb,A(AV)= - 1 4 . 3 fps)

System capacitya I -- 1 5 2 4 4 .4

4 8 .2

Offloadedb I 2 0 . 7 1 5 2 2 3 .7

Maran I - - I 4 8 .2I I' 2028 .0 pounds of fuel and 3218 .4 pounds of oxidizer.bF'uel offload of 20 . 7 pounds o minimize malfunction penalty.

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required 6056-fps AV is established by the nominal insertion into a 9 - by 45-nautical-mile orbit. In addition, a 54-fps AV is provided fo r two contingencies. A 40-fps AVis provided fo r the first contingency, which is a switchover f ro m PGNCS to AGS for in-se rti ng fro m an off-nominal traject ory caused by a malfunctioning PGNCS. A 14-fpsAV is provided for the second contingency, in which the thrust -to-weight rat io is re -duced in an abort fro m a touchdown situation wherein the LM ascent stage is heavierthan the nominal ascent -stage lift-off weight. Some weight is nominally off-loaded onthe lunar surf ace . Also, 19 pounds of propellant is allotted for contingency engine-valve malfunction, as in the descent requirements. The allowance for dispers ions isdetermined from the Monte Carlo analysis. A s can be-seen in table 11, the AV andpropellant r equireme nts are satisfied with a positive marg in of 48 pounds.

REAL-T IME ANALY S ISDuring the real- time situation, monitoring of the spac ecra ft sys te ms and of thetrajectory is performed continually both on board by the cr ew and on the ground by theflight cont roll ers. The real-time monitoring determines whether the mission is to becontinued or aborted, as established by mission techniques pr io r to flight. The re al -

time situation for the Apollo 11 descent and ascent is described in the following section.D e s c e n t O r b i t In s e r t i o n

The DO1 maneuver is performed on the fa rs id e of the moon at a position in theorbit 180" prior to the PDI and is , therefo re, executed and monitored sole ly by thecrew. Of major concern during the burn is the per formance of the PGNCS and the DPS.The DO1 maneuver is essentially a retrograd e burn to reduce orbit altitude fro m approx-imately 60 nautical miles to 50 000 feet for the PDI and requi res a velocity reduction of75 fps. This reduction is accomplished by throttling the DPS to 10-percent thrust for15 seconds (center-of-gravity trimming) and to 40-percent th rus t for 13 seconds. Anpriorby rarthe m;with tlthe b u

systergrouncsiblemust 1guidardesce

16

overburn of 1 2 fps (or 3 seconds) would cause the LM to be on an impacting trajectoryPDI. Thus, the DO1 is monitored by the cre w with the AGS during the burn and?-rate tracking ;with the rendezvous r ad ar (RR) immediately after the burn. Ifeuver is unsat isfactory, an immediate rendezvous with the CSM is performedAGS. For Apqllo 11, this maneuver w a s nominal. Down-range resi dua ls aft er

1 were 0.4 fps. ,I

P owered Descenth e powered descen t is a complex maneuver which is demanding on both cr ew andperformances. 1 Therefore, as much monitoring as possible is performed on theto reduce crew activities and to u s e sophisticated computing techniques not pos-board. Obviously, however, time-cri tical fai lures and near -sur face operations1 monitored on board by the crew fo r immediate action. Pertinent aspec ts ofe, propulsion, and rea l- tim e monitoring of flight dynamics during the poweredare given as follows.

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The PGNCS monitoring. - To determ i degraded performance of the PGNlground flight con tro lle rs continually comp the LM velocity components ,cornpithe PGNCS with those computed by the AGS and with those determined on the grcthrough Manned Space Flight Network (MSFN) tracking. That is, a two-out-of-tvoting comparison logic is used to de NCS o r the AGS isgrading. Limit o r redlines for 'veloc he PGNCS and th e Mcomputations and between the PGNCS and the AGS c ns/ar e establishedthe mission, based on the ability to abort on the PGNCS to a safe (30 OOOlfoot pcorbit.

In rea l time, the Apollo 11 PGNCS and AGS performance kas close to nonhowever, a la rg e velocity difference in the radia l directio n of 18 fps (limit linewas detected at PDI and remained constant well into the burn. iThis e r r o r did ncate a sys tem s performance problem, but rath er an initialization e r r o r in downposition. This effect is illustrated geometrically in figure 16. 1 The PGNCS posRE and velocity ?E estimates are used to initiate the MSFN dowered-flight prThe MSFN directly senses the actual velocity VA at the ;actuad position RA, bihaving been initialized by the PGNCS state, the MSFN applies )A at RE. ThLflight-path-angle error A is introducedby a down-range position e r r o r and shows

I 1

2

A 2

A

Y ,I

S, theted byind.reele-IF Ntefore:ilune)

.nal;t indi-range:ion)cesser.

t 35 fps)

- 9

i, a

Aup as a radial velocity difference AVDIFF.The magnitude of the velocity d ifference in-dicates that t he Apollo 11 LM down-rangeposition was in e r ro r by approximately3 nautical miles at PDI and throughout thepowered descent to landing. The rea sonfor the down-range navigatattributed to se ver al sma llThese inputs w ere f ro m uncoupled RCS at-titude maneuvers and cooling sy stem vent-ing not accounted for in the prediction of thenavigated state at PDI.ing computation ta sk s: navigation, guidance, displays, ra da r data processing, andauxiliary tasks. If the computer becomes overloaded o r falls behind in accomplishingthese tasks, an ala rm is issued to inform the crew and the flight contr oller s, and pri-or i t ies are established so that the more important tas ks a re accomplished first. Thisalarm system is termed "computer restart protection. 'I During real time, because ofan improperly defined inte rface, a continuous signal was issued to the LGC from theRR through coupling da ta units. Thes e signals caused the LGC to count pulses contin-ually in an attempt to slew the RR until a computation time interval w a s exceeded. Asa result, the ala rm w a s displayed and computation priorities were executed by the com-puter. The ala rm was quickly interpreted, and flight-contro l monitoring indicated that

the spacecraft state in coasting flight. Center of moonFigure 16. - Effect of position er r o r onvelocity comparison.

The LM guidance computer (LGC) ls o monitors the speed at which it is perform-

17


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guidance and navigation functions were being perf ormed properly; thus, the descent w a scontinued. In sp it e of the initial position e r ro r and the RR inputs, the PGNCS per form edexcellently during the Apollo 11 powered descent.

diver@tory, (isfaccfor thivideseven vHowevGTC ttweension rabortthe AgnearlyPGNC

The DPS and PGNCS interface . - To determine in real tim e if the DPS is providingsufficient thrust to achieve the guidance tar gets , the.flight co ntro lle rs monitor a plot ofguidance thrust command (GTC) as a function of hor izonta l velocity, as s hu re 17. Nominally, the GTC de cr ea se s almos t parabolically fr om an initi160 percent of design thrust to the throttleable level of 57 percent, approximately2 minutes (horizontal velocity being 1400 ps ) before high gate (horizontal velocity being500 ps). If the DPS produces off-nominal high t h r u s t , horizo ntal velocity is being re -duced more rapidly than desired to reach high-gate conditions. Therefore, the GTCdrops to 57 percent earlier with a higher-than-nominal velocity to guide to th e desiredposition and velocity tar get s. This early throttledown r es ul ts in propellant inefficiency.If the DPS produces off-nominal low thrust , horizontal velocity is not being reducedrapidly enough. The refo re, the GTC drops to 57 percent later at a lower velocity toguide to the de sired position and velocity. This later throttledown results in incpropellant efficiency (i.e., longer operation a t maximum thrus t). However, iftledown occ urs pr ior to high gate (p rogram switch from P63 t o P64), the targets willnot be satisfied, and the result ing trajec tory may not be satisfactory from the stand-point of visibility. In fact, fo r extremely low thru st, the guidanceisolution or the GTCcan diverge (fig. 17); as TGO becomessmall, the guidance call s for more and morethrust in or der to achieve its targets. This-

2 0 0 .ic e can result in an unsafe traiec- leoL '\. T\..Ie from which an abort!cannot be sat-ily per form ed. The 2;minute b iasttle recov ery before high gate pro-tfficient marginfor 3a low thrustch propellant valve ma,Hunction.r, the flight contr olle rs monitor theassure satisfacdory in ter face be-PS and PGNCS operation. A mis-e was established that, called for anwed on the GTC'divergence. During110 11 anding, the DPS th rus t w a siominal (fig. 17); thus,, no DPS andinterface problems were encountered.

PD

----- 3 0 disprsions9350

Horizontal velocity, fps\Figure 17. - Guidance -thrust commandas a function of hor izonta l velocity.

estimpersicdescewheremissiscentestabl

concealtituc18

I

Ihe LR and PGNCS int4rface. - Normally, the LR update of the PGNCS alt itudee is expected toioccuriby crew input at an altitude of 39 000 +_ 5000 eet (3a dis-). Withou t LR altitude updating, syst em and navigation e r r o r s are such that thecannot be safely completed. In fact, it is unsafe to tr y to achiev e high gatehe crew can visually assess the approach without altit ude updating. Thus, ai rule for rea l-time oqeration w a s established that called for aborting the de-:a PGNCS-estimated altitude of 10 000 feet, if altitude updating had not been;bed.i addition to the concern for the time that init ial altitude updating o ccu rs is thei for the amount of alt itude updating (i. e. , the difference between PGNCS and LRdeterminations Ah). If the L M is actually higher than the PGNCS estima te, the

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Lunar Module Descent and Ascent