lunar landing symposium

8/8/2019 Lunar Landing Symposium

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NASA-S-66-5944 JV L 1APOLLO LAUNCH OPPORTUNITIES IN 1969 THAT

PROVIDE A 1-3-5 DAY LAUNCH WINDOW FORORBITER B SITES

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P i w e P A

APOLLO LAUNCH OPPORTUNITIES IN 1968 THAT PROVIDE A1-3-5 DAY LAUNCH WINDOW FOR ORBITER B SITESN A S A - S - 6 6 . 5 2 0 6 IU N

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LAUNCH I I I I: IEST HRS

TIME, JAN FE E MAR APR MAY JUNE

"[I20 j :5 . . Donald C. CheathamAssistant Chief f o rEngineering and DevelopmentGuidance and Control DivisionFloyd V. BennettAssistant ChiefB e o r e t i c a l Mechanics BranchGuidance and Control Division. .


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INTROrnCTIONSTRA'BCX CONSIDEEATTONS

Spacecraft Systems

Guidance and Control System

bang Radar SystemD J Window Systemk s c e n t Propulsion System

Mission Ian- Position RequirementPOWERED DESCENT DESIGN

Wing PhaseObjectives and Constraints

Igni t ion LogicGuidance with Limited ! throt t leLanding Radar U p d a t i n g

Delta V &@etDescent Guidance MonitoringSurmnzry of Braking Phase

Final Approach PhaseObjectives and ConstraintsDetermination of Hi-gateParameter h d e o f f sRedesignation Footprint

Ianding Point Designator Utilization


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APOLLO LUNAR MODULF: LANDING STRATEGY

Delta V Buagetf lmnnary of Final Approach Phase

Landing Phase

Objectives and Constraints

Nom in al TkajectoryDelta V Budget

LUNAR LANDING TouCKCKlWN CONTROLObjective and Constraints - Modes of control

Sequence of events

Descent engine shut-off

Manua control of landing veloci t iesAutomatic control of landing veloci t i es

ABORT AFTER TOUCBDOWNSUMMARY

1.0 INTRODUCTION

The landing of the Lunar Module (IN) upon the surface ofth e moon w i l l be th e climax of th e Apollo mission, althoughthe importance of the r etur n phases'is not to be de-emphasized.The IN landing approach w i l l be the fi r st t ime tha t the com-ple t e LM system w i l l have been operated in th e luna r environ-ment. This a l s o w i l l be m a n ' s i n i t i a l f ac e-to-face encounterwith the exact nature of the te rrai n i n the landing area andof the problems of vis ib il it y as they may affec t the ab ili tyto l and the LM; although, these aspects of the landing w i l lbe simulated many times in fixed-based simulators and partialpre fli ght simula tors. These simulations are extremelyimportant i n the preparations fo r the mission; but onlyafte r the mission i s completed w i l l it be known how adequatethe simulations have been.

Considering the entire LM descent after separation from theComand Module i n lunar orb it, a theoretical landing maneuvercould consist of a Hohmann transfer impulse on the back sideof the moon with a d e l t a V, o r change in vel ocity, of lo9f t / s e c , followed 180 later by an impulsive velocity changeof about 5622 f t / s e c as th e LM approaches the lunar surface,as i l l u s t r a t e d ih figure 1. The fl ig ht path angle in thefi na l portio n of t he approach would be zero degrees. Sucha theoretical approach would require infinite t h ru s t - t o -weight ra ti o by th e descent engine. This, of course, i san impossible and impractical approach. A f i n i t e t h ru st-to-weight ra ti o of th e descent engine must be used and theapproach path must account f or luna r te rr ain v ariatio ns anduncer tainties i n the guidance system. Since lunar te rr ai nvariat ions of as much as + 20,000 f t . could be expected, and,a l so , uncertaint ies in t h e value of the lunar referenceradiu s, coupled with guidance dispe rsio ns, could add anotherl5,OOO ft . to t he uncertainty, a conservative safe value of5O,OOO f t . was chosen as a pericynthion alti tude . From aperformance standpoint, the choice of 5O,OOO f t . as opposedt o e i t he r 40,000 or 60,000 f t . was quite arb itra ry becausethe dif ference from the standpoint of fue l requirements wa svery sl ight , as indicated in f igure 2. The ini t i al thrust-to-weight of the LM descent engine w i l l be about three- tenths.Combining this thrust-to-weight with a perigee al t i tude of50,000 f t . l eads t o t h e descent prof i l e , as shown i n figure 3 .The separ ation and Hohmann tra ns fe r maneuver req uires sl ig ht ly


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l ess de l t a V due to the pericynthion alti tude increase.The powered descent portion approaching the landing area,however, requires a d e l t a V of 5925 f t / s e c , which i s a con-siderable increase over the infini te thrust requirement. Ascaled t rajectory prof i le of th is theoret ical LM powereddescent i s shown i n f igur e 4, indicat ing that the ent i redescent takes approximately 220 n. m. The LM velocity andat t i tude i s shown periodically along the flight profile.Th i s traj ecto ry has the predominant ch aracte ristics of alow flat profile terminating with a fl ig ht path angle of

about 9 degrees. An obvious feature i s th at the crew, con-sidering the location of the LM window, never have theopportunity t o see where they ar e going. They can lookei ther di rect ly up, or , i f th e LM i s rotated about i t sthrust axis, can look down at the surface, but they ar enever able to see in th e dir ecti on they are going. Ifthe crew i s t o perform any assessment of the landing areaor out-the-window sa fe ty of f li gh t during the approach, iti s obvious that the la tt er por tion of the traj ecto ry mustbe shaped so t ha t a differe nt at t i tu de of the LM can beused during the approach. Shaping the tra jec tor y awaJfrom the fuel optimum approach w i l l result i n a penaltyi n fue l requirements. Both the amount of time the creww i l l require to assess the landing area, and the rangefrom which the landing area can be adequately assessedmust be traded off again st th e amount of f ue l involved

i n the penalty of the shaping. It soon becomes obvious. t ha t a st rategy i s needed that w i l l trade off the system

capabi l i t ies of the spacecraft and the crew capabilitiesagainst the unknowns of the lunar environment encounteredduring the descent from the o rb it, in order to insure th atproper u ti li za ti on of the onboard systems can be made t ogre ate st advantage. The development of thi s strateg y,then, i s the subject of th is paper.

1.0 STWEGY CONSIDERATIONSThe LM landing strategy can be defined as the science andand art of s pacecraft mission planning exercised t o meetthe lunar environmental problems under advantageous condi-t ions. In order to plan strategy, the object ives, theproblems to be faced, and the cha rac ter isti c performance

of availa ble systems need t o be we ll known. As indicatedin f i gure 5, the objectives of the M landing planningstrategy are to anticipate the lunar environmental pro-blems and to plan th e lan ding approach so tha t the com-bined spacecraft systems, including the crew, w i l l mosteffec tively improve the pr obabili ty of a ttaining a safelanding. The major fac tor s tha t must be considered i nthis strategy are the problems brought about by the

or bi ta l mechanics of t he landing maneuver, the limita tion sof the spacecraft systems (including limitations i n fue lcapacity and payload capabilit y), and the con straints ofthe lunar environment (ircluding t e r ra in uncer ta in t ies ,visibility, and determination of sui table landing posi t ions),The or bi ta l mechanics aspect s have been discussed i n thepreceeding secti on. The lunar environmental constraintsw i l l be discussed i n a subsequent sec tio n. The remainderof t h i s sec t ion i s concerned with descriptions of the space-cra ft systems and the mission landing posit ion requirements.

Although all of the LM systems are important t o at ta in thelunar landing, those affecting the st rategy are (a ) theguidance and con tro l system, ( b) the landing radar , (c) thespace craf t window, and (d) t he descent propuls ion system.

Spacecraft Systems

Guidance and control system - The guidance and control systemi s important to the landing strateg y i n t h a t it has a di rec tef fe ct upon the a re a over which th e lan din g may be accomplishedand on the problems of lan din g at a desir ed point. The func-t ional description and accuracies of this system have beendiscussed i n a preceeding paper. The ef fe ct of the guidance,navigation, and con trol system of the LM on the landing beginswith navigation in the lunar or bi t. The accuracy of thisnavigation, whether performed by the onboard system o r byth e Manned Space Fli gh t Network, determine s th e unc er ta int iesat th e start of th e powered des cent . Assuming t ha t t heguidance system w i l l be updated by landing radar t o eli-minate the a lti tud e dispersions, th e landing dispersionsw i l l be a funct ion of the i ni t i al condi tion uncertaint iesbrought about from lunar or bi t navigation coupled with th einert ial system drift during t he powered descent. A summaryof the guidance system capability for attaining a givenlanding point on the moon i s presented i n figure 6a andthe asso ciated assumptions i n fig ure 6b. Both th e MSFNand the spacecraft onboard navigation in l unar orb i t a reconsidered. The Apollo system spe cif ica tion of a landingCEP of 3000 f t . i s met i n eith er case when the ine rt ia lsystem performs within specification.

The 30- landing dispersion ellipses are shown in figure 7for cases where the lunar orbit navigation wa s done by theMSFN and also onboard the CSM. The ell ips es are quite

similar with the major axis for t he MSFN case being slightlyshorter and the minor axis for the MSFN being sl igh t ly longerthan t ha t for t he case u t i l i z ing CSM onboard naviga tion. Aspe cial case i n which the downrange dist ance was allowed t o


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be unconstrained i s also shown on figure 7 - I n t h i s casethe downrange or major axis of the ellipse i s primarily afunct ion of the thrust uncertaint ies of th e fixed- throt t leposition of the descent engine that wi ll be discussed sub-sequently. The crossrange axis i s equal to tha t of the 3Celli pse s fo r guidance to a spec ific point and i s determinedby the method of lunar orb it nav igation.

Landing radar system - The con trol of the LM during thedescent t o the surface can be provided automaticallythrough steering commands generated by the guidance systemand a l s o manually by the crew by inputs through an attitudecont rolle r. The primary control system sta bili zat ion uti liz esdi gi ta l aut opil ot mode of th e guidance computer. Figure 8shows the att itud e thruste r f ir in g combinations to createcontrolmom ents. The engines are loc ate d on an axes systemrotated about the LM descent engine thrust area 45' from thespace craft axes. They are operated as control couples forthree-axis at t i tude control . As can be seen in figure 8,two pairs of control couples are available for each axis.The method of providing tran sla tio nal cont rol while i n th ehovering condition i s t o tilt the spacecraft by means ofthe atti tud e contro l system. This produces a l a t e r a lcomponent of accele ration from the descent engine thrus ti n the desired d irection which i s stopped by retur ning t over t ica l and reversed by t i l t in g i n the opposite di rect ion.

During the descent the attitude control system is a l so coupledt o a slow moving gimbal actuator system of the descent engineto enable a means of trimming the descent engine thrust direc-t i on so t h a t it passes through the LM cente r of gravit y. Thetrimming system reduces undesirable torques from the descentengine i n order t o conserve RCS propellant. The LH landingradar system i s important in landing strateg y. As indicatede a r l i e r , it i s used to eliminate the guidance system alti-tude dispersions and, also, the uncertainties of knowing th ealti tud e from the lunar surface pri or to beginning the descent.The LM landing radar i s a &-beam dopple system with the beamconfig uratio n shown i n figure 9. The c en ter beam measuresthe a ltit ude , and the other three beams measure the th ree

components of velocity. Two posi t ions of the landing radarI antenna provide both altitude and velocity measurements overI a wide range of spacecraft attitu des. The fi r s t antenna

posi t ion i s t i l t ed back from the t h r u s t axis by approximatelyforty- three degrees so tha t the al t i tud e beam w i l l be nearlyve rti ca l during the ear ly portions of the descent and, hence,w i l l s t i l l provide accurate altitude information. As th e LMapproaches the lan ding maneuver, th e antenna i s physicallyswitched t o the second positi on making the al tit ud e beam

p a ra l l e l t o t h e X - a x i s of th e LM The landing radar w i l l

begin t o provide a lt itu de measurements a t an approximatea l t i t ude of 40,000 f t . These altitude measurements w i l lbe used to update th e i n e r t i a l system starting at analtitude of about 25,000 f t . The radar velo city updateswill begin at approximately 15,000 ft . The landing radaraccuracy i s given i n figure 10 .LM window system - The LM window, alth ough perh aps not nor-mally considered a system, is a very important part of thelanding strategy because it is through this window that the

crew must observe the landing are a to confirm the adequacyof the surface f o r touchdown. The phy sic al con fig ura tio nof the LM window i s shown i n f igur e ll. T h i s photographwa s made from with in t he LM cockpit showi-ng the left hand,or t he command pi lo t' s, window. The window i s t riangulari n shape and skewed so t h a t it provides maximum viewingangles f or t he landing approach maneuver. Although thewindow is not large in size, the pi lot 's eye posi tion i snormally very c lose t o th e window so that the angular limitsprovided are qu ite wide. These angular limits are displayedi n f i g ure l.2, showing th e limits as viewed from the commander'sdesign eye pos ition . The plo t shows the azimuth and elevationvariations of possible viewing limits referenced from a pointwhere the p il ot would be looking dead ahead, with respe ct toLM body axes (p aral le l to the Z-body axis) , for the zero point.It i s pogsible for the pi lot t o see downward at an angle ofabout 65 from the normal eye position and to the l e f t sideby approximately 800. I f the pi lo t moves his head eit herclos er t o th e window, o r furt her back, these limita tions changes l ight ly .

The guidance system i s coupled with the window system throughgr id markings so that the pilot can observe the intended land-ing area by al igning his l ine-of-sight wi th the gr id markingaccording t o information displa yed from the guidance system.Figure ll in addition t o showing t&window system, shows th elocation of the Display and Keyboard, which among other thingsprovide di gi ta l readout information from the guidance system.The procedures fo r uti liz ing the se integr ated systems fo rlanding site designation and redesignation w i l l be discussed

la t er i n t h i s paper .

Descent propulsion system - The descent engine i s an extremelyimportant system to the de sign of the LM descent strategy.In iti all y, the descent engine was capable of being thr ott ledover a range from 10 t o 1. Design considerations, however,have made it necessary t o limit t he t hro t t l e capabi l it y t oth at shown in figure 13. This figure shows tha t a t the start


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of Dowered f l igh t , there i s an upper fixed position of thethrottle which would nominally provide about 9700 l b . oft hrus t . A s long as t h e t h ro t t l e i s maintained i n this fixedposit ion, t hr ust magnitude w i l l vary according to th e nominalsol id l ine. A t th e start of the powered fl igh t, t here i sexpected to be approximately 1percent uncertainty i n thethrus t at t h i s f i xed- throttle setting. The uncertainty growsup t o 2 percent after approximately 300 seconds of f ixed-th ro ttl e usage. The descent engine i s always t h ro t t l eabl e ,

I n the region of 6300 l b . of t hrus t , t o approximately 1050-1b. of th ru st . The change from a ful ly throt t leable enginei n the upper region of the t hrus t lev el to a fixed- throt t leposition aff ects the guidance procedures during the i n it ia lpowered descent, as w i l l be explained later.

Mission Landing Position Requirement

Important stra tegy considerations axe the types of require-ments that are placed on the landing position, as indicatedi n figure 1 4 . The first consideration i s a requirement t oland at any suitable point within a specified =ea, w i t h theimplication that th e ar ea could be quite large. Obviously,i f th e asea i s lar ge enough, the requirements on the guidancesystem would be diminished con side rabl y. The second type ofrequirement i s that of landing at any suit able point within

a reasonably s m a l l area, constrained in si ze primarily bythe guidance dispersions. T h i s would, of c o K se , dic t a t etha t t he s i ze of th e area chosen w i l l be compatible with thecap abil itie s of the guidance and navigation system. Thethird consideration i s that of landing at a prespecifiedpoint, such as landing w i t h 100 f t . of the position of asurveyor spacecraft, or perhaps another type of spacec raft.It i s obvious that this latter consideration imposes thegre ates t requirements on the strategy and also the guidancesystem, and would require some means of establishing contactwith the intended landing pos itio n during the approach. Thepresent strategy i s primarily based upon the second consider-at ion, that of landing in areas of the size compatible withthe guidance system dispersions. If, however, th e landin g

area. can be increase d i n si ze t o the point th at downrangeposi t ion control i s not of primary importance, the associated

strategy i s not great ly di fferent than that f or the require-ment assumed because the trajectory shaping requirementswould be the same for the terminal portion of the trajectory'.The subsequ.ent discussions of this paper w i l l be basedprimarily upon a landing area size compatible with guidancesystem dispersio ns.

3. PCWEXED DESCENT DESIGNAfter consideration of all the t rade off 's t hat could beident i fied as worthy of consider-ation during th e LM powereddescent, a three-phase t ra jector y design logic was chosen.The logic of this three-phase trajectory design w i l l b ediscussed in the subsequent section s, but, the general logi ci s indicated i n figure 1 5 - The f i rs t phase following powereddescent ini t iat ion at 50,OOG f t . i s termed the Braking Phase.This phase i s terminated at what i s cal led a Hi-gate position.

The second phase i s termed the F in al Approach Phase, and i sterminated at what i s cal led the Lo-gate position, the startof the Landing Phase. The to ta l tra jec tor y covers on th eorder of 250 n.m. The logic of the braking phase i s designedfor the efficien t reduct ion of veloci ty. That i s , since therei s no necessi ty for pi lot vis ibi l i ty of the landing area i nth is phase, the a ttitu des can be chosen so that the spacecraftwould have efficient utilization of descent engine thrust forreducing velo city . During th e fin al approach phase, thet ra j ec tory i s shaped t o a l l ow an attitude from which thepilot can visually acquire and assess the landing site. Anadditional requirement met'by thi s phase i s t o provide thepi lot wi th a view of the terr ain at such a time that he canconfi rm the f l ig ht safety of the t raje ctory prior to committ ingt o a landing. The land ing phase i s flown very much as a VTOLtype o f air cr af t would be flown on the ear th t o allow the

pi lo t vernier control of the posi t ion and veloci t ies at touch-down. The a tt it u de chosen i s flown so as t o provide t he crewwith v i s ib i l i t y fo r a detailed assessment of the landing site.The scaled profile of the design descent trajectory i s showni n f i gu re 16 a.) and b) , and includes an indicati on of th espacecraft at t i tude at various milestones along the trajectory.The f in al approach and landing phases together cover only about2 per cent of the tot al tra jecto ry range, although the timespent within these phases w i l l be about 30 per cent of theto ta l time. The following sectio ns w i l l di scuss i n de t a i l

t he de l t a V budget for the descent." the logic of the design of the three phases and w i l l summarize_ . .

Eraking PhaseObjectives and constraints - The objec tive of th e brakingphase, as indicated i n figure 17, i s to provide efficie ntreduction of the horizon tal velocity ex isting at t he i n i t i a-

ti on of the powered descent. During most of th is phase, t heal t i tude i s high enough so tha t the pilo t does not have t oworry about the t e r r a i n variations, and he can conduct thereduct ion i n veloci ty at at t i tu des that al low great efficiency.The major constraint of this t rajectory phase i s l imi tat ionsimposed by the fixed- throt t le-posi t ion thrust of the descentengine. It i s des i rab l e t o use the maximum t h rus t of t hedescent engine as long as possible i n order t o provide effici entut i l iz at ion of the fuel . There is , however, an in i t ia l segment


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of the powered descent which i s flown at reduced throt t le t oinsure tha t t he descent engine gimbal tr im mechanism hasnulled out of trim moments due t o ce nter-of-gravity offse ts.

Igniti on Logic - The log ic fo r ig niti ng the descent enginefor ini t i at io n of the braking phase i s as fol lows. Fir st ,th e LM stat e ( posi t ion and veloci ty) i s integr ated ahead intime. Next, t he guidance problem fo r the braking phase i ssolved, b ut not implemented, continu ously with the advancedLM s t a t e s as in it ia l conditions. When the guidance soluti onrequires the level of thrust equal t o the expected thrust ofthe fixed- throt t le posi t ion, see figure 18, that solut ion i schosen fo r init ia ti on of braking. Finall y, when the LMreaches the posi t ion and veloci ty s tat e that yielded theproper thrust solution, the guidance computer sends theengine on signal to the descent propulsion solut ion. Inorder to prevent large moments due to c. g. of fs et, theengine i s igni ted at the low 10 percent level, instead ofmaximum th ru st . This le ve l i s held f or some 28 seconds totrim the engine gimbal through the c.g. before increasingthrus t t o t he maximm, or fixed- throt t le, set t ing. Thislow level of thrus ting is accounted for i n t he i gni ti onlogic.

Guidance with Limited Throttle - The general approach oftlie braking phase, from the stan dpoint of t he guidancesystem, is t o ut il iz e the same type of guidance equationsthat are appropriate for the throttled phases which follow.

Thus, modifications i n the tar geti ng are requ ired to allowfor the u t i l i za t i on of t he f i xed- throttle position duringthis phase. It i s s t i l l desired to vary the st ate vectorof the LM from i t s value at t h e start of powered descent

to t he s t a t e speci f i ed at t he h i-gate posFtion of the tra- ject ory . The guidance equa tions wouldnormally determinethe thrus t l eve l or accelerat ion leve l and at t i t ude requiredin order to make an effi cien t change i n the state . Prior

knowledge of the in it ia l thrust- to-weight of the descentengine allows choice of in i t i a l conditions and the guidanceequat ions to be ut i l i zed in such a way as t o s e l e c t a timeto go for the ent i r e phase that w i l l use the approximatethrust-to-weight of th e upper limit of the descent engine.

In actual operation, the LM system during t hi s phase w i l lrespond to comands of a tti tud e change, but as long as t heguidance system i s cal l i ng for a thrust above 6300 l b . , th edescent engine w i l l remain i n i t s fixed or upper limitposition. If the thru st va riat ion of the descent engineat th is fixe d thr ot tle positio n were known exactly, th etrajec tory could be preplanned to obtain th e desired h i -gate state vector. In view of the uncertaint ies of thedescent engine, however, th e tra jec tor y must be plannedso that the guidance system will b eg in t o c a l l f o r t h ru s t

leve ls i n the region i n which the descent engine can bethrottled (below 6300 l b s . ) p r i o r t o reaching Hi-gateposition. This i s to provide control over the v e l o c i t i e swhen the Hi-gate position i s reached. The log ic of t h i sguidance scheme i s shown i n f igure 19. The figure showst he prof i l e of t he t ra j ec tor i es as a function of range,and also a profile of the descent engine thrust, both thenominal value and that Commanded by the guidance system asa func tion of range. The nominal thrust- to-weight case i sshown f irs t , and the t ra jectory i s .essen t ial ly preplanned

by f lyi ng backward from the h i-gate pos i t ion , f i r s t o f all,using a t h rus t i n t he t hro t t l eabl e range to go back for aperiod of t ime; the period of t i m e being determined by the-poss ible magnitude of t he un certainty of the descent engine.This, i n effec t , determines the fict i t i ous targ et that can

be used i n the guidance system i n the first port ion of th etrajec tory. The fi ct iti ou s tar ge t is'based upon the nominalthrust profile when the descent engine i s i n t he f i xed- thrustpositi on. The logic of the guidance i s perhaps best explainedby comparing the actual value of thrust with that commanded bythe guidance system, even though in th e upper thr us t regionthe descent engine i s not responding t o these commands.

Initially, the guidance system i s targeted to a f i c t i t i o u starget upstream of the hi-gate state. The nominal thrust-to-weight variation follows the solid line, and the guidance

system computes the commanded variation of thrust- to-weightshown on the figure. A t an intermediate position, t heguidance t argeti ng i s switched from that of the fi ct i t io ust arge t t o t ha t of t he hi-gate tar get . The discon tinuityseen i n th e commanded posi tio n has no effect on the system,since, i n this region, th e descent engine throt t le i s no tresponding to the guidance system. I f the thrust- to-weightdoes remain nominal, th e commanded thr us t-to-weight m a g n i-tude will gradually decrease mtil it i s within the regioni n which the descent engine can be throt tled . This w i l lnominally occur at t he f i c t i t i ous t a rge t pos i t ion . The

guidance system then has a number of seconds, pri or t o thehi -gate position, t o match both the veloc ity and the positi ondesi red at hi -gat e. From hi -gat e on, t h e commanded th ru stwill be at o r below the m a x i m u m i n t he t hro t t l eabl e range.

Figure 20 il lu st ra te s the t hr ust pr of ile s (commanded andactual) for low and high thrust- to-weight rat i os. In thecase of the low thrust-to-weight ra ti o where the actualvalue of the thru st i s below th at of t he expected nominal,

it i s seen tha t the i n i t i a l commanded thru st has t he sametype of variation as the nominal, pri or t o the switchoverpoint ; but , af ter the switchover point , there i s a delay intime and range i n get ting down t o th e region where the

commanded thr us t reaches the thr ot tl ea bl e region. Thispoint , then, i s only a few seconds prior to hi-gate. The


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extreme low thrust-to-weight, then , would be tha t i n whichthe commanded thrust would reach the t b z o t t l e a b l e regionthrust exactly at th e time the hi-gate position was reached.For the case where the thrust- to-weight i s higher than nomi-nal, the commanded thrust w i l l reach the throt t leable posi t iona number of seconds pri or t o tha t fo r nominal thr ust . Thisallows a much longer time t o affe ct the desi red veloci ty con-di t i on a t t he h i-gate position. This, however, means tha t

the region prior t o hi-gate i s being flown at a much lowerthrust-to-weight rat io for a longe r period of time than would

be desirable from a standpoint of fue l effic iency. This i sthe case tha t involves the greate st penalty in fuel. Figure2 1 shows the delta V penalty vari&c,ion due to fixed- thrustunce rtain ties. The left-hand scale indicates the del ta Vpenalty, the horizontal scale shows the bia s time of thefict i t iou s target back from the hi-gate target , and the right -hand scale i s the thrust- to-weight uncertainty expressed i n fpercentages. The figur e indica tes tha t the 2 2 percent un-cer tai nty of the descent engine w i l l require a bias time ofapproximately 65 seconds and w i l l invoke a bias de l t a V pen-al ty on the order of 45 f t / s e c . In ef fec t , t he 45 f t / s e c . off u e l i s the penalty paid for reducing th e landing dispersionsfrom that associated with the range-free type of guidance, t othat in which a desired posi t ion at hi -gate i s reached. Themagnitude of additional variation i n the landing point tha twould be associated with range-free type of guidance i sessentially the percentage uncertainty thrust-to-weight valuetimes the tot al range trave l. For the case of5 2 percentaverage thrust uncertainty and a nominal range of 250 n. m. ,which can be eliminated at th e-cost of 45 f t / s e c . of fuelt h i s results i n approximately + 5 n. m. of range uncertaintypenalty.

Landing Radar Updating - The. effe ct of landing radar (LR)updating on the guidance commands i s important from thestandpoint of eliminating alt itu de uncerta inties, and theresul ting changes in attit ude and thro tt le required by thechange in soluti on of the guidance equations. The effe ctof landing radar update i s a continuing ef fec t throughoutthe t rajectory once the in i t ia l update al t i tud e i s reached;and, therefore, some aspects of the following discussions

w i l l touch on the fi na l approach phase as well as the brakingphase.The altitude update i s i n i t i a t e d at 25,000 f t . , as determinedby the primary guidance system, and i s continued a t each two-second int er va l fo r the remainder of the approach. Velocityupdates are initiated at about 15,000 f t . , when the velocityi s reduced t o about 1550 f t / s e c . The velocity is updated a

a single component at a t ime, in two-second intervals (6seconds for a complete update). The alti tu de updating i scontinued along with t'ne velo city components. After each

complete (3components) velocit y updating, an alt itu de up-date only i s performed, then the velo city updating i s con-t inued. The weighting facto rs for LR al t i tude and veloci tyupdates a r e i l l us t ra t ed i n f igure 22 as l inear funct ions ofth e parameter being updated. These are lin ea r approximationst o optimum weighting based upon lea st-squares estimation.

The guidance commands f o r an ideal descent (no ini t i al con-di t i on er rors , no 7Mo- errors, no LR errors , no t e r ra in var ia-t ions, no DPS uncer tainties ) ar e shown in Zigure 23 . Thetraje ctor y presented in the figure i s not the nominal designt ra j ec tory , bu t i s adecpate t o i l l u s t r a t e t h e e f f e c t s o f landing radar update. This partic ular tra ject ory has a hi -gate al t i tude of 6100 f t . and a throt t le period of 80 se c .p r i o r t o h i-gate. The pitc h pro file exhibits a slope dis-cont inui ty at t he f i c t i t i ous t a rge t poin t (TF) for t hro t t l i ngthe engine, as shown in par t (b) of the f igure .

At the hi-gate target point (HG) , the pitch angle undergoes

near constant (about 35O of the vert ica l ). A t the low-gatet a rg e t (TLG, about 500 f t . a l t i t ud e) , t he a t t i t ude begins t ochange (nearly l inear) t o sat i sfy the ne= ver t i ca l a t t i t udedesi red just prior t o the vert ica l descent targe t (TVD, about100 f t . al t i t udej , 10' off the vert ica l . The prof i le i sterminated at t h i s poin t .

The same traje cto ry has been analyzed for cases with i ni ti alcondition errors, descent engine -thrust uncertaint ies, IMUerrors, landing radar errors and a t yp i ca l t e r ra in prof i l eapproaching the landing si te . The terr ain pr of ile used i sshown in fig ure 24 and i s applicable for an approach to as i t e at 020'N lat i tude and 12'30'E longitude. Both aproDer1y scaled profiie and an expanded altitude scale pro-f i l e a re shown.

An example effe ct of the terra in, in it ia l condition and systemerrors i s shown in figure 25 . In addi ti on to t he e f fec t of the t err ain the other i ni t i al predominent err or included was

an al tit ude uncertainty of about minus 1600 fe et. This casei s considered somewhat extreme in th at t he al ti tu de uncerta iz tyo f -1600 feet i s about a 3c~magnitudei f CSM landmark typesightings have been maae on the landing si te and in a di rec t ivesuch the ter rai n effects are addi t ive -with the inert ial systemalt itu de uncertainty tending to accentuate the pi tch angle andthru st variat ions from the idea l case. The time hist ori es o f

pitch angle and thrust magnitude are presented i n f igure 25and include the id eal case t o provide a basis for comparison.

I>_lLe rapid pi tchup t o t he const ant a t t i t ude des i red for f i na l


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The pitc h angle vari es by sl igh tly more than 10 degrees ata maximum pr io r t o hi -gate and is about equivalent afterhi -gate. The th ru st le ve l shows genera lly the same le ve lof command. The pit ch angle de via tion s are of concern becauseof possible effect upon landing radar operation and becauseof increased expenditure of descent engine propellants.

In the event tha t no landmark sightings near the landing si teaxe performed in lunar orb it, large un certa inties (up t o 10,000f t . on 3 r b a s i s ) in the braking al t i tude can exist . Invest i-gations of the abi lit y of the LR t o update these large altitudeuncertain ties have indicated tha t 100 fps of add itiona l de lta Vi s requi red. Furthermore, thr ot tl e commands above 60 percentand large attitude deviations (up to 70') occur i n some in-stances in the th rot tle down region prior to hi-gate. Furtherinvestigation of th is problem i s proceeding.

Delta V Budget - The nominal fue l expenditure dur ing the brakin gphase i s 5206 f t / s e c . To this an addi t ional 1 5 f t / s e c . i s addedto account for po ssible mission changes th at would rai se t he CSMal t i tude 10 n.m. For the random thrust uncertainties of thedescent engine a 36random fuel expenditure of 2 20 f t / s e c . i sbudgeted. In addition , analy sis has shown th at navigatio nuncertaintie s i n alti tud e, although eventually eliminated bythe landing radar, w i l l change fu el consumption by about 60f t / s e c . for a 3000 f t . uncertainty. TO account f o r t h i s , a3 6 random fue l expenditure of

i60 f t / s e c . has been allottedon the fu el budget.

Descent Guidance Monitoring - An important func tion of the crewduring the braking phase i s t o monitor t he performance of theguidance system onboard. This i s done by checking the solutionof the primary guidance system with the solu tion of pos itionand velocity obtained from the abort guidance system. As indi-cated in figure 26, t h i s i s accomplished by periodic differ -encing of the primary and abort guidance solutions of altitude,al t i tude rate, and lat er al veloci t ies. The al t i tud e rate para-meter i s perhaps the most si gni fic ant parameter t o monitorbecause th is i s the one that c a l lead to a t ra j ec tory tha tviolates the fl ight safety zonsiderations. Analysis has shown,however, that it w i l l take greater than the extremes of 36 per-formance of the abort and primary guidance s oluti ons t o lea d

to an unsafe t rajectory prior to the hi-gate pos ition . Becauseth e Mznned Space Flight Network w i l l be very effect ive i nmeasuring the alti tude r ate of the spacecraft, it also w i l lbe very effec tiv e in providing an independent vote in theevent that onboard differencing indicates the possibilityof a guidance fail ure . The tot al procedures fo r th is guidancemonitoring are s t i l l in the formative stages and are cu rrentlybeing in vest igat ed i n simulations conducted by the MannedSpacecraft Center.

Summary of Braking Phase - The braking phase, lasting about$50 seconds, covers some 243 naut ical miles during which thevelocity i s reduced from 5500 f t l s e c . to approximately 600f t / s e c . , and the altitude from 50,000 feet t o about 9,000feet. The at titu de during the phase is normally such tha tthe thrust vector i s close t o being aligned with the f lig htpath angle. In this at t i t ude, the pi lot i s not able to lookin the di rect ion of the intended . landing area. In the f i r s tportion of t M s phase, the LM could assume any desired rollattitude about the X o r thr us t axis . Mission planning will

determine i f t he i n i t i a l a t t i t ude w i l l allow the crew t o lookdown on the l una r surf ace t o check the progres s over theterr ain. As the LM approaches the position at which landingradar w i l l begin operat ing, the r ol l at t i tude w i l l be suchthat the windows w i l l be oriented away from the s urface i norder to provide a more favorable attitude for the landingradar operation and to prepare for th e pitch-up maneuver atthe hi-gate position that w i l l allow a view forvrard t o t helanding area.

Final Approach Phase

Objectives and Constraints - The final approach phase isperhaps the most important phase, from the standpoint ofthe st rategy. It i s primari ly in t his phase that the tra - jectory i s shaped at a cost of fuel, in order to provide

the crew with vi si bi li ty of the landing area. In thi s phase,the crew begin s t o be confr onted with some of the pos sibl eunknowns of t he lunar environment, such as the po ss ibi lit yof reduced visib ili ty. The objectives of the fin al approachphase are enumerated i n f i gure 27. The fi rs t object ive i s t oprovide the crew with out-the-window vi si bi li ty , and toprovide adequate time t o ass ess th e landing are a. The secondis t o provide the crew with an opportunity to as sess thef l i gh t saf e ty of t he t ra j ec tory before committing th e contin-,uat ion of the landing. And thir dly , to provid e a re l a t i ve lystabl e viewing platform i n order to b est accomplish the firstand second objectives. In other words, lnaneuvering should bekept to a minimum. The pr-ry constraints on the strategyi n this pbase are again the desi re t o keep the fue l expendi turet o a mininnun and the limit ation of the LM window. In theevent that the ascent engine m st be used f o r abort during thisapproach to the surface, the difference i n thrust-to-weightbetween the descent and ascent engines must also be consideredas a cons trai nt. The asce nt engines thrust-to-weight i n i t i a l l yis only abaut one -half of that of the descent engine i n t h i sphase. The al ti tu de loss during ver t i ca l v e l o c i t y a u l l i n g asa f'unction of nominal t raject ory al t i tude and veloci ty must beincluded in the consideration for a safe staged abort. Theother constraints that must be considered are the problems of


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It has Seen assumed that a maximum capability of designating3000 f t domrange w i l l be required and this provision of fu e lis a l l o t b e d for redesignation a t 5OOO f t of altitude. Approx-imately 45 f t / s e c of fuel is requi red f o r this redesignationcapability. Figure 37 shows the footprint available f o rthi s f ue l allotment.

The LMpilo t does not have the opportunity t o see th e foot-pri nt as viewed here, but inste ad from the perspectiveprovided by the approach fl ig ht pat h angle. The pi lo t view

from the hi-gate al t i tude i s indicated in figure 38. Wri n #this phase, the space craft is pitched 3 c k a p pr o xb a te l y 40 ,t h s , the horizon i s very near the - 40 ele vation depressionangle. The landing si te i s a t approx-tely 55' depressionor about 10' above the lower limit of th e window. Forreference purposes a 3000 f t c i r c l e has been drawn about thelanding position and the landing footpr int asso ciated with.

a del ta V of 100 f t / s e c i s shown.Landing Point Designator - The pilot will know where t o lookto f ind th e intended landing area, o r th e area which the@dance system is tak ing him, by informa tion earning fromthe guidance system disp lay and keyboard (DSKY). This infor-mation will be in the form of a di gi t al readout that allowshim to locate the correct grid number on th e window, comonlycalled the landing point designator (LPD). After properalinement of the grid, the pi lot merely .has to look beyondth e number corresponding to th e DSKY readout to fin d whereon the lunar surface the automatic system is guiding thespace craft . The proposed gr id config uratio n f o r the landingpoint -designator is shown in f i gure 39.The process of landing point designation and redesignation isi l l us t ra t ed in f i gure 40. The guidance system always believesthat it is following the correct path to the landing si te. Ithas the capabi l i ty a t any time t o determine the proper lookangle or l ine-of-sight t o the intended landing site. Becauseof orb i tal navigat ion errors and also drifts of t he i ner t i a lsystem &ring the powered descent, the a ctua l positio n ofthe spacecraft will not be the correct position. Thus, ift he p i lo t looks along the calculated line-of- sight he wouldse e an area different from that of the desired landing area.Should the desired landing area appear i n another portion ofthe window, then th e pilo t, by taking a measurement of th eangle formed by the line-of-sight readout from the guidancesystem and the new line-of-sight ( to the desi red point),can input the change in line-of-sight into the guidance computer.

will be a cooperative ta sk between the pi lo t and the copilo twhere the copilot will read the DSKY and ca l l ou t t o t hepi lo t the numbers corresponding to the landing point designator.The pilot will then orient his l ine-of-sight so that he canlook beyond the prop er number on the landing poin t de signato rand see where the guidance system is taking him. If he i s no tsat isf ied with th is position, then he can instr uct changes inthe guidance system by incrementing his attitude hand controller.=ring t h i s port i on of th e approach, t he guidance system i sflying the spacecraft automatically so that t he p i lo t ' s a t t i t ude

hand controller i s not effect ive in lnaking attitude changes.With each increment that the pilot makes in moving the handcontrol ler in a pitching motion, there i s an inst ruct ion tothe guidance system to change the landing point by th e equiv-alent of a half-degree of elevation viewing angle. I a t e r a lchanges i n th e land ing posit ion would be =de by incrementingthe hand control ler to the side in a motion that would normallycreate ro lling motion of the spacecraft. Ehch increment ofa hand control ler in th is di rect ion causes a 2 l ine-of- sightchange late ra l ly t o the landing area. When the guidance systemreceives these discrete inst ruct ions it recalculates theposi t ion of the desired landing area and comrnands t h e p i t c ho r r o l l a t t i t u d e in combination with a t h ro t t l e commndrequired to reach the desi red posi t ion. This resul ts i n at r a n s i e n t response from the spacecraft until the new attitudeand throt t le set t ing commnds ar e responded to. After thet ransient has se ttl ed out, the copilo t would n o r m l l y readth e DSKY again and inform the pi lo t w h a t new number t o l oo kf o r t o f i nd the a e s i re d landing area. The pi lo t would thenorie nt himself t o looka t this number and check to se e i fhis i ns truc t ions t o t he guidmce system had been fu ll y corre ct.E not, some refinement i n landing s it e selec tion would thenbe made.

The response of the spacecraft t o redesignations of landingposi t ion is important. For example, if the new si te se lect edi s f i r t h e r downrange, the spacecraft will pi t ch c loser t oth e verticral and reduct ion in thr ot t le wi l l be made so thatthe new position will be more closely centered in t h e p i l o t ' swindow. If, however, th e s i t e chosen i s short of t he or ig ina llanding sit e, then the spacecr aft would have to pitc h t a c kand increase throt t le in order to slow down and obtain thenew desired position. These atti tud e motions aff ect the l i n e -of-sight and become important because of the danger of losingsight of the target . Some t m i c a l responses to changes i nthe landing point ar e shown in figure 43. The variation ofthe line-of-sight to the landing s i t e (looking angle) wi th

time from hi-gate i s shown f o r the nominal case, a redesignation-


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w r y of Final Approach Phase - The final approach phasecovers about 5 1 /2 naut ical miles during wbich t he a l t i t udei s decreased from 9 0 0 f t t o 5 0 t , and the velocity from600 f t / s e c t o 50 f t / s e c . The time required n o m l l y w i l l beabout lo5 seconds during which.time t h e p i l o t will b v e acontinuous view of the landing area. It i s during this timethat assessments of the landing area w i l l be made, andrequired redesignations of the landing position to morefavorable landing terrain will be accomplished.

The Landing Phase

Objectives and Constraints - The basic purpose o f the landingphase is to provide a port ion of fl i ght a t low veloci t ies andat pi t ch a t t i tudes c lose t o t he ve r t i ca l so t ha t t he p i lo tcan provide vernier control of th e touchdown maneuver, andalso to have the opportunity for detailed assessment of th eare a pri or t o the touchdown. In order to accomplish thi s,the t rajectory is mrther shaped after the final approachphase. The guidanc e system i s targeted so that the designconstraints of the lo-gate position are m e t , bu t the actualtarget point w i l l be a t or near the position where the verticaldescent begins. The final approach phase and the landingphase are then combined with regard t o th e mn n e r in whichthe guidance system i s targeted. The +arget ing design mmldsat - i sfy the constraints of both the terminal portion of th efinal approach phase and the landing phase by propersele ctio n of the targetin g parameters. There will be asmooth transition from the extreme pitch-back at t i tude withassociated with the final approach phase and the near verticalatt itu de of th e landing phase.

In t he final approach phase, th e t r a j e c t o r y w a s shaped i norder to pitch the att itud e more toward the ver tica l so t ha tthe approach conditions would allow the pi lo t t o view thelanding si te . The resul t ing pi tc h at t i tude , approxTmately40' back from the vertical i s , however, s t i l l quite extremef o r approaching the lunar surface at low altitudes, hence,it i s necessary to provide ad dition al shaping in order toeffec t a more nearly vert ica l at t i tude a t the termination o fthe t ota l descent. Figure 44 shows a comparison of th enominal attit ude s fo r those two phases. ,The objectives andconstraints of the landing phase design ar e presented infigure 45. The first objective i s to allow the crew t o makethe detailed assessment, and a f i n a l se l ec t ion , of the exactlanding point. In order t o accomplish this, there will besome fle xib ili ty in the prop ellant budget t o allow other thana rigid following of the design traje ctory . This leads toobjective m b e r two, i n which it is desired to allow some

maneuvering capability and adjustment of the landing point.The constraints are familiar ones including the f u e l ut i l i za-t ion, the physical l i d t a t i o n s of th e window, and i n turn ,the l ight i ng and associated vis ibi l i ty of the surface, t h evis ibi l i t y associated with the l ight ing, the actual terrain,and the po ss ibi lit y of blowing d u s t maneuvering within th edesired at t i tude limits in order to re ta in th e advantages of afa ir ly stabl e platform, and last, what is termed the stagedabor t limiti ng boundary. This boundary define s the ci rcustancesunder which an abort maneuver cannot be performed without theascent stage hit tin g the surface. This curve is based upon acombination of ver tica l velocities, altitu des, and the p i l o t-abort-staging system reaction time.

Nominal Trajectory - The variables that are avai lable t o t r y tos a t is f y a l l of these constraints and objectives include variationsin the approach flig ht path and the vel oci ties involved, theat tit ud e of the spacecraf t, and the act ual touchdown controlprocedures. The lacding phase p ro f i l e which has resulted fromalmost & years of simulating the maneuver i s i l l u s t r a t e d i nfigure 46 The lo-gate point is a t an alti tud e of approximately500 ft ., a t a position about 1200 f t back from the intendedlanding spot. The landing phase fl ig ht path i s a continuationof the final approach phase flight path so t ha t t here is nodiscont inui ty a t the lo-gate position. A t t he s t a r t of t h i sphase, the horizontal velocity is approximately 50 f t / s e c andthe ver t i ca l veloc&ty i s 1 5 f t / s e c . The pitch att itud e i snominally 10 t o 11 throughout this phase, but rigid adherencet o t h i s p i t c h a t ti t u d e i s not a requirement. The ef fe ct ofthe pi tch at t i tude i s to gradual ly reduce the veloci t ies asthe f l ig ht path i s followed i n order t o reach the desi redposi t ion at an al t i tude of 103 f t from which a vert ical descentcan be made. Modif icati on of this trajectory can be accomplishedsimply by modifying the profile of pi t ch a t t i t ude i n order t oeffec t a landing at sl ight l y di fferent points than that associatedwith the nominal descent path. No actual hover position i sshown in the approach porfil e because th e v er tic al v e b c i t yo r desc ent ra te nominally does not come to zero. The approachi s a continu ous maneuver in which forward an d l a t e r a l ve loc i t i eswould be zeroed a t approximately the 100 f t al t i tude posi t ionand the descent velo city allowed to continue a t approximately5 f t / s e c . This allows a very expeditious type of landing, however,i f a hover condition i s desired near the 100 f t alti tude mark.It i s a very simple matter f or the pi lo t t o eff ect such a hovermaneuver. The onl y disa dvan tage of th e hover maneuver is th eexpenditure of fue l. The to ta l maneuver from the lo-gate positionw i l l normally take approximately 89 seconds. If flown accordingto the profi le, the descent propel lant ut i l i zed w i l l be equiva-


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lent to about 390 f t / s e c of characterist ic veloci ty. Duringthe landing approach, th e pi lo t has good visi bi li ty of th elanding posi tion u nt i l j us t before t he f i na l ver t i ca l descentphase. Figure 46 al so shows a nominal sequence of p il ot viewsof a 100 f t radius c i rcu l ar a rea- around the landing point.However, even during the ve rt ic al descent, the a re a immediatelyin fron t of and to the side of the exact landing position w i l lbe vis ibl e. The IM fro nt landing pad is vi s ib l e t o t he p i lo t .In addi t ion t o being able to observe the intezded landingsi te , th e pi lo t has ample view of much of the lunar surface

around h i m so t h a t i f t h e o r ig i n a l s i t e is not suitable hecan deviate to the othe r landing position; provided t ha t thenew landing pos iti on is sb t a i n a b l e with the fue l avai lable.The basic system design w i l l allow the e nt ir e maneuver to beconducted aut oma tic all y. However, th e IM handling qualitiesmake it a s at isfactory vehicle f or the p i lo t t o control manual ly.The satis fac tor y nature of the LN manual control handlingqualities has been demonstrated by fixed base simulation and byfli gh t simulation at the Flight Research Center using the Lunar

Landing Research Vehicle and the Langley Research Center usingthe Lunar Landing Research Fa ci lit y. Simulations have showntha t here should be no problems involved i f the p il ot de cidesto take over from manual con trol a t any time during the terminalportion of t he f in al approach phase o r th e landing phase.

Much concern has been generated with regard t o th e problem ofvi si bi lit y during the landing approach. This fac tor has led toa constraint upon the sun angle at the landing site, as w i l lbe discussed by the paper on Site Selection. In the event thatthe pilot has some misgivings about the area on which he desirest o land, the landing phase can be fl ex ib le enough to accommodatea dog-leg type maneuver that w i l l give the pilot improved view-ing perspective of the intended landing position . Manual contro lof t hi s maneuver should pre sen t no problem and could be executedat the opt ion of t he p i lo t . A t the present time, t ra j ec toryi s not planned f or an approach i n order t o Icaintain simplicityof tra jec tor y design, because of th e expected ease in which themaneuver could be accomplished manually should the need be present.Should, however, th e dog-leg be id entif ied as a requirement f or

an automatic approach, it w i l l be incorporated.

A profi le of the al t i tud e and al t i tu de ra te of th e landing phasei s shown i n figure 47. The altitude rate is gradually decreasedt o a value of about 5 f t / s e c a t t h e lC3 f t al t i tude posi t ion forve rt ic al descent. The descent ra te of 5 f t / s e c i s maintained a tthi s point in order+to expedite the landing. A t approximately50 f t of a l t i t u d e (- 10 ft ) , the descent ra te would be decreasedto the design touchdown velocity of 3s f t / s e c . It is not necessary

f o r this to be done a t exact ly 50 f t so that uncertaint ies i nthe altitu de of the order of 5 t o 10 f t would not s i g n i f i c a n t l yaff ect the approach design. The value of 3%f t j sec descentr a t e i s then maintained a l l the way un til contact with thesurface is effected and procedures in itia ted f or cutoff of t h edescent engine. The curve l a b e l e a staged-abor t boundary showni n figure 47 is applicable t o the sit uatio n in which the descentengine has to be cut off and the vehicle staged t o abor t onthe ascent engine. It i s obvious that this boundary must beviol a ted pr ior t o e f fec t ing a normal landing on the surface.However, with the cur re nt design, th is boundary i s avoidedunt i l t he p i lo t is ready t o commit hiplself t o a landing so t ha tit

is only in the region of below 100 f t t h a t he i s i n violat ionof t he boundary.

Delta V Budget - A sunrmary of the landing phase fu el budget isgiven in f igure 48 The budget reflects allowances for severalpossib le contingencies. For example, the pi lo t may wish t o pr o-ceed to t he landing s it e and spend some time inspect ing it beforehe fi nal ly descends t o the surface. This would require t ha t the

spacecraft hes ita te during the approach, and the penalty involvedi s the amount of f ue l expended. A period of 15 seconds of hovert i m e w i l l cost about 80 f t j s e c of fu el equivalent. There is alsothe possibility. that the performance of the landin g rad ar may bedoubtful, i n which case the s pacec raft crew might want t o hoverin order t o visu ally observe and nu l l out t h e veloci t ies. Ithas been found by means of fl ig ht t es ts i n a helicopter, th atveloc ities can be nulled in this manner within 1 t / s e c a f t e rless than 1 5 seconds of hover time (another 80 f t / s e c of f u e lexpenditure). It would be possible t o update the in er ti alsystem i n this manner and allow the s pacecr aft t o proceed andl a n d on the surface with degraded landing radar performanceduring the fi na l port ion of the descent . I f there ar e e r r o r sin the radar vert i cal veloci ty, there w i l l be a d i r e c t e f f e c tupon the time req uir ed t o complete descent and a random 2 65f t / s e c of equivalent fu el has b e e n .a l l o t t e d in the fu el budget.Another descent engine fuel contingency that must be accountedf o r is the possible variat ions i n the pi lot control techniqueincluding the deviations from the planned flight profile the

pi lo t might employ. Simulation experience has ind icat ed a needfor an average addition of &I f t / s e c of fuel and a random * 103f t j s e c . It i s noteworthy that only 30 seconds of hover time hasbeen budgeted and that f o r spec ifica lly designated purposes.


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Fuel Budget Summary

A summary of the to ta l -24 descent fu el budget i s given infigure 49. The budget i s divided into that required by thebaseline trajectory requirement totaling 6582 f t / s e c , anditems, described as Contingencies, tot ali ng 353 f t / s e c meanrequirement with an additional 2 143 f t j s e c random requirement.This leads to a t ot al 7053. The inclusion of the RSS randomcontingencies as a f ue l requirement i s considered a conserva-t iv e approach in th at each of the random congingencies couldlead to a fu el savings as well as a f e u l expenditure. Thepre sen t tankage would provide up to 7332 ft/sec of fuel orabout 282 f t / s e c more than the budget. Thus, the pos si bi lit yof additi onal landing fl ex ib il it y can be provided by fue l tanks,or i n the inte re st of weight savings, some off-loading of fuelcan be considered. The additio n fl ex ib il it y i s equivalent toa hover time of about one minute o r to add ition al downrangelanding redesignation capa bility of almost 2D,OM f e e t f o r aredesignation at . 8,300 f t a lt it ud e .The f ue l budget s m a r y i s presented i n figur e 50b as a How-Goes-It plot of the expenditure of fue l both i n equivalentchar acte risti c velocity and pounds as a function of time andevents during the descent, The solid li ne give the baselinetrajectory and r e s u l t s i n a fuel remaining of 778 f t / s e c a ttouchdown. Adding the uti li za ti on of a l l of the budgeted con-tingency mean values of fu el i s represented by the dashed line..When these contingencies are u tilize d th e time basi s of the plo tw i l l be incorrect, par ticul arly f o r the time between Lo-Gate andTouchdown. The t o t a l time could extend t o as much as 12minutes(735 seconds) i n the event .that a l l of the contingency fue l wereutili zed f or hovering over the landing site .

3.0 LUNAR LANDING TOUCHDOWN CONTROL, AUTCMATIC AND MANUALPerhaps the most important singl e operation in the lunar landingmission is the actual touchdown maneuver. It i s during thismaneuver that the unce'rtanities of the lunar surface become areal problem. A recommended procedure fo r c ontrolling the approachhas been developed. This procedure, developed par tl y throughsimulation, involves reaching a pos itio n at about 100 f t abovethe landing si te and descending ver tic ally t o the lunar surface,as previously described. Diring the vert ica l descent , thelat era l veloci t ies are nul led and the vert ica l veloci ty control ledto a prescribed value until the descent engine is cu t off jus tpri or to touchdown. The procedures fo r ef fec tin g descent engineshutdown w i l l be discussed i n detai l .

. i

There are two control modes by which the landing operation canbe performed, as indicat ed i n f i gure 51. The f ir s t i s completelyautomatic. In this mode, whil e th e pi l o t may have used th elanding point designator t o se le ct th e touchdown point, he i snot active i n the actual contr ol loop. The second mode i s manual,bu t i s aided 5y automatic control loops, t ha t is , t he p i lo thas taken over direc t control but he has sta bil iza tion loopst o provide favorable control response. In addition, t he manualmode normally w i l l be used i n conjunction with a rate-of-descent,comand mode to f urth er aid the p il ot in control of the touch-down ve lo ci ti es . Within the manual landi ng mode, the p il othas two options; (1)land visu ally , which would require that

th er e be no vi su al obs cura tion a s might come from dus t o r lunarl ight in g constraints, or (2 ) because of such obscurations hewould control the landing through reference t o f l i g h t in s t ru-ments. Because of the expected good handling qua lit ie s of theIM the manual vis ual mode should be very similar t o fl ig ht ofa VTOL air cra ft here on earth. No landing at t i tu de o r veloci tycont rol problem i s ant icip ated and the c ontrol should be with-

i n one foot per second late ra l veloci t ies. Manual-instrumentmode of control does have sources of err or, however, t ha t maydegrade cont ro l and those t h a t have been considered include thefollowing: cont ro l system response, landing radar vel oci tymeasurement, landi ng ra da r al ti tu de measurement, IMU accelero-meter bias, IMU misalignment, d isp lay system fo r manual only,the pi lot, for manual only, and the center-of-gravity (c .g . )position. Several of these parameters are listed in figure 5 1as being of prime importance.

I n considering the cont rol of t he landing, emphasis has beenplaced on the method of timing of s hut tin g of f the descentengine. Because of pos sib le unsymmetrical nozzle f ai lu re duet o shock ingestion and a desire to l imi t erosion of the landingsurface, an operatin g -constraint of having the descent engineoff at touchdown has been accepted. Probable erro rs i n alti tu deinformation from either the inertial system or from the landing

radar preclude the use of this information f o r the engine cut-off func tion, even though th e accur acy may be of the or der offiv e feet , because of the deleterious effect on touchdown v e r t i -ca i veloc ities . The need fo r an accurate, disc rete indicationof the proper altitude to cut the descent engine off l ed t o t headoption of probes extending beneath the landing pads rigged t ocause a l igh t i n the cockpi t t o turn on upon probe contact witht he lunar surf ace. The ligh t-on signal informs the p i l o t t h a t

the proper al t i tude -ha s been reached fo r engine cutof f. Theprobe length must be determined from a consideration of delaytimes i n pilo t response, descent engine shutoff valve closures,

and tail-of f and the nominal descent velocit ies. The sequence

of events i s shown i n f i gure 52.


18/39

The variation of descent ra te a t touchdown as a function ofdescent r ate at probe contact i s shown i n f igure 53, and includesthe ef fec t of pi lo t reac tion time. The curves are representativeof a 53-inch foo t probe being used, coupled with a 0.25 secondto ta l engine shutoff delay time. This engine delay time includes

that time required for the elect ronic signal to be generated,the shutoff valves to close, and the thru st tail -o f f t o b e

es sen tia lly completed. The heavy dashed li ne on the ch artgoing up on a 45 degree angle indicated a combination of descentra te a t probe contact, plus system delay and pi lo t reactiontimes, th at would cause the engine t o still be on a t touch-down. If the desi red fin al rat e of descent has been achieved,

up t o 1. 0 second pi lo t delay time can be tole rat ed an d s t i l lhave the descent engine o ff a t touchdown. A s shown in figure 53,the actual toucMown veloci ty i s just s l i ght ly more than thedescent rat e a t probe contact, or about four feet per second.Faster reaction time would increase the fi na l touchdown velo-city, but not beyond present landing gear impact limit. Ifmanual contro l allowed a sl ig ht ly higher than desired fi na ldescent rate, and radar error s a t the time of fin al update alsoallowed a s lig htl y higher descent r ate , thes e compounded in-creases might yield descent rates on the order of 5 t o 6 f t / s e c .These increased rates coupled with the 0.6 second reaction timewould mean not meeting the cr it er ia of having the descent engineoff a t touchdown. One solut ion f o r th is sit uat ion would be toextend the probes t o allow more leeway i n pi lo t r eaction time.HrJWeVer, the advantages of longer probes m u s t be traded of fagainst a probable decrease i n reliability and an increased pro-bab ilit y of touching down with greate r than acceptable vert ica lveloci t ies. A simulation study of this maneuver with the pilotcutoff of the descent engine showed that pi lo t reaction timesaveraged about a.3 seconds, a s shown i n figure 54.

Pilot-in- the-loop and automatic control simulation studies havebeen conducted of the landing cont ro l maneuver. The p i l o t - i n -the-loop studies were made using a s i a a t e d LM cockpit includinga l l the control actuators (at t i tude, thr ot t le and descent engine

cuto ff) . The simulation included the major sources of systemerro rs, such as platform misalignment, accelerometer bias, instru -ment display resolu tion, center -of-gravity offs ets, and landingradar errors. The landing radar er ror s are a prime fact or in th etouchdown contr ol process and the models assumed f or t he a nal ys isar e shown i n figure 55. The specification performance of t helanding radar calls f o r each of the three components of veloci tyt o be measured within 1. 5 f t / s e c on a 3 W basis. Currect pre-dict ions are that this specificat ion w i l l be met i n lat er al andforward directions and bettered by 3/4 f t / s e c vert ical ly. Fora conservative analysis, the pre dicted performance has beendegraded by a factor o r two.

I

The simulation resu lts of landing velocity manual control withspec ifica tion performance by th e l a n d i n g radar are shown i n fig-ure' 56. The dashed lines indicate the present design criteriafo r the landing gear. The 0.9, 0 . s and 0 .99 probability con-tours are shown and are well within the design envelope. Theeffec t of changing the length of th e l m d i n g probes i s t o a d ju s tthe ve r t i ca l ve loc i ty b i as veloc'i ty approximately 1 t / s e c pe rfoot change. i n probe length.

The effect of landing rada r performance upon the landing vel oci tyenvelope is shown i n figure 57. The 0.99 probability contoursare shown f o r the cases of no radar errors, specificat ion per-

formance, pre dict ed performance, and degraded (pr edic ted ) per-formance. The resulting contours show the almost direct depend-ence of touchdown velocity error upon the landing radar velocityperformance.

The comparative results between automatic and manual control ofth e landing touchdown veloc itie s are shown in f igure 58. The 0 . 9contours show that automatic control results i n lower touchdownveloc ities, but th e difference i s much les s pronounced fo r th edegraded radar performance as compared with the predicted radar

performance. The fi gu re does not, of course, re fle ct the advant-age tha t manual control prov ides i n closer select ion of the actualtouchdown position in th e event tha t the te rr ain is not uniformlysat isfactory.

Additional analysis of these same res ults fo r the control per-

formance for a t t i tude and at t i tude rates indicated tha t controlwithin the present cr i ter ia of 6 degrees and 2 degrees per sec-ond can be expected on a 3 a p r o b a b i l i t y .

4.0 ABORT AFICER 'XXTCHDOWNAlthough analysis and simulation tests indicate a high probabilityth at th e la ndi ng touchdown maneuver w i l l be within the l a n f i n ggear design cri teria, there i s still a n in t e r e s t i n t h e a b i l i t yt o abor t should the landing dynamics become unstable. The ab il it yt o a b or t w i l l be a function of when the need for the abort i srecognized, the t ime required to in it ia te abort, the time involvedin separation of the ascent stage, the thr ust buildup time of th eascent stage, the at t i tud e and the at t i tude ra te a t separation,and the control power and con trol ra te l imitat ions of the ascent

stage.

A t staging, the control power of the ascent stage is about 35deg/sec2 fo r pitch and roll at ti tu de maneuvers. Under emergencymanual control where the -pi lo t def l ec t s his at t i tude hand c o n t ro l l e r


19/39

hard-over, there i s no at tit ud e ra te limitat ion. Normal manualcontrol commands are l imited to a O / s e c and automatic controll imited to 10/sec in p i tch and S0/sec i n r o l l . These attitudera te limita tions are important from the standpoint of determining

how quickly the ascent stage a ttit ude can be returned t o thever t ica l in the event of an impending t ipover.An an aly sis was made of th e boundary of ov er-turn conditionsfrom which a succ ess ful staged abo rt could be made. The res ul tsof this analysis are shown in figure 59. Two boundaries areshown; one fo r emergency manual at ti tu de co ntr ol which requ ire sthe p i lo t t o put his hand controller hard over and the other

fo r a r a t e limit consistent with automatic roll response ( j o / se c ) .Both boundaries apply to t he conditions under which an abor tact ion must be recognized as being require d. The boundar iesal low a tot al of 1. 4 seconds for the time required f or the pi lo tto actuate the abort control, the staging to take place, andthe ascent thrust t o bui ld up t o 9 percent of rated thru st.In addi t ion t o the boundaries, there is al so a l i ne i n d i c a t l n gthe neutral sta bi l i ty boundary o r t he se t s of condition underwhich the spacecraft would just reach the tipover balance point

of about 40 degrees. The curve lab eled Landing Gear DesignEnvelope Maximum Enegry applies to the improbable, i f not . im-poss ible, case where the landing was made a t the corner of th e

veloci ty cr i ter ia envelope 7 f t / s e c vert ical and 4 f t / s e c horizon-ta l, and a l l of th e energy was converted t o rotat ion al motion.It is , therefore, highly improbable that conditions w i l l be

encountered that l ie to the right of th is curve.

For the emergency manual contr ol, th e boundary indic ate d anabort can be made a t an al tit ud e of about 63 degrees i f ther a t e is not greater than 13 degjsec. This condition would takemore than 4 seconds t o develop af ter the in i t ia l contact wi thth e lunar surface. For the other extreme of at titu de ra te limit(5O/sec) applicable only t o automatic r o l l at t i tude control, theboundary i s reduced about 13 degrees i n at t i tude .The pilot w i l l have indicatio n of a ttit ude from kiis winsowview and from the attitude instrument display (FDAI). Bothof these are considered adequate sources of attitude informationin the event that the spacecraft passes a 43 degree deviationfrom the vertical and an abort becomes necessary.

Considering th e improbabi l i ty of landing contact that wouldr e s u l t i n an unstable post-landing attitude and the probabilitytha t even in such an event the pi lo t could in it ia te a sa fe abort,the re does not appear t o be a requirement f or an automatic aborti n i t i a t i on .

The preceedin g se cti on s have describe d and explained th e [email protected] f th e IM descent st rategy and the resul t ing t raje ctory desi=-.From th e p i l o t ' s standpoint there are a number of judgmentsand decisions that w i l l have to be made i n the period fromH i Gate t o Lo Gate t o touchaown. It is believed that thestrategy allows a logical sequence of events and decisions andadequate time fo r the pi lot funct ion. This w i l l be part lyconfirmed or adjustments made through extensive simulationswith the I M Mission Simulators. The fi na l confirmation w i l l ,of course, be the resu l ts of the f i r s t I M landing approach.In order t o aid in the understanding of the l ogic and proposed

sequence of decisions, a logic-flow chart has been preparedtha t i s applicable from the H i :Gate pos ition to landing tauch-down. These ch ar ts are presente d i n figures 63a) and b) forthe information and use of persons interested in detai ledexamination of th e log ic and in const ructing the crew loadin g

time lines. & t a i l s of these logic flow charts w i l l not bediscussed fur ther in thi s paper.

6.0 f l l " ~ ~ yA IM descent s tra teg y has been presented which i s designed t otake advantage of the 161 system and the IM crew in order th atth e IM w i l l continually be i n an advantageous posi tion t o corn-ple te the lunar landing. The three phases traje ctory i s designedto maintain fu el expenditure efficiency, except i n those regions

of the t raj ect ory where such factors as pil ot assessment ofthe landing a rea requ ire a judicious compromise of f ue l effici ency.

The lunar lan ding st ra teg y has considered all ident ified problemswhich might adversely affec t the lunar landing and the resultingdesign calls for a fuel expenditure budget of 7350 ft / sec o fcharacterist ic veloci ty. This budget. i s approximately 282 f t j s e cl e s s than the current tank capacity of the I" This margin i sconsidered ample f or dealing with pres entl y unforeseen problemswhich may be ide ntifi ed pri or to the lunar landing.

N A S A - S - 6 6 - 6 0 2 5 MAY


20/39

Questions and Answers

LUNAR EXCURSION MODULF DESCENTSpeaker: Donald C. Cheatham

1. Mr. Kelly - Probabi l i ty plots of landing velocity showconstant vert ical veloci ty f o r a l l probab i l i t ies whenhorizontal velocity i s zero; is this correct?

ANSWER - Mr. Kelly and Mr. Cheatham discussed the dataaf te r the meeting and resolved thei r differe nces on the

presentation form.

THEORETICAL LM DESCENT

IMPULSIVE A V

SEPARATION AND

TRANSFER

A V c = 109 fT/SEC

lMPULSlVE POWEREDTERMINATION

A V c = 5622 FT/SEC

FIGURE 1TOTAL AVC = 5731E R MIN A L Y Oo

N A S A - S - 6 6 - 5 0 3 9 JUNVARIATION OF POWERED-DESCENTCHARACTERISTIC VELOCITY WITH

THRUST-TO-WEIGHT RATIO

[Eo (PERICYNITION ALTITUDE, FT)

INITIALTHRUST-TO-WEIGHT

RATIO

FIGURE 2

II

I

0 1 I I56 5a 60 6 2 6 4CHARACTERISTIC VELOCITY, FPS


21/39

N A S A -5 - 6 6 -6 5 0 4 JU N NASA.S .65-1606


22/39

LM LANDING ACCURACYAFTER THREE O RBITS(C0 NT)ASSUMPTIONS & ERROR MODELSI1"l

0 LANDING SITE AT 0" LATITUDE AN D 0"LONGITUDE0 MSFN UPDATE PRIOR TO LUNAR O R B I T INSERTION

0 TW O LANDMARKS WITH T H R E E S l G H T l N G S PE RLANDMARK PER PASS

0 LM SEPARATION FROM CSM O N THIRD ORBIT,PLATFORM ALINEMENT AT 15 MINUTESBEF O RE A MANEUVER

ACCEL BIAS

ALINEMENTACCURACY (ACT)

LANDMARKACCURACY

NAS A-5 -66 -6522 JUN

LM LANDlE

,0017 FT /SE C2 ~ .06 DE G

ATTITUDE

CONTROL OF LM

NOTE:

IN DECENT THRUST CONFIG-URATION MAIN ENGINE

GIMBAL IS EMPLOYED FOR

TRIMMING THE PITCH AND

YAW MOMENT DU E TOCENTER OF GRAVITY

SHIFTS

.06 DE G GYRO DRIFT .03 D E G / H RFIGURE 8

7500 FTFIGURE 68

4G 3 0 UNCERTAINTY ELLIPSEAFTER THREE ORBITS

LANDING SITE 0" LAT 0" L O N G

ORBIT BY LMBY MSFN

N A V I G A T I O N I N N A V I G A T I O N I N

ORBIT BY LMLUNAR ORBIT

BY MSFN

15,000 \5000 0 2 5 0 0 15,000

NASA. S. 6 6 . 5 0 5 0 J U N E 1LANDING RADAR

BEAM CONFIGURATION ANDANTENNA TILT ANGLES

FIGURE 9

+ xPOSITION NO. 2

ANTENNA TILT = O oSITION NO. 1NN A TILT = 43

L 7500r 7500

I I I10.000 o-2500

-5000GUIDANCE TARGETED FOR 7 5 0 0

N O N C O N S T R A I N E D R A N G E D " FIGURE 7

N A S A 5 . 66 . 6479 JUN N A S A - S - 6 4 5 0 d 5 JUNE IO VIEWING LIMITS FROM


23/39

LM LANDING RADAR (30)SPECIFICATION ACCURACY

I ACCURACYALTITUDE, FT

RANGE TO

SURFACE

5 . 20 0

1.5% OR 1 .5 FPS1 . 5 % t 5 F T2000 - 25,0001 . 5 % OR 1. 5 FPS1 . 5 % i 5 F TZ OO . 20001.5% OR 1.5 FPS1 . 5 % t 5 FT

25,000 . 40,000 N/A2%

I

I

'"A, 'ZA I2.0% OR 1. 5 FPS3.5% OR 3. 5 FP S

2.0% OR 2.0 FPS

N/A

1M W I N D O W VIEWING LIMITS FROMC O M M A N D E R ' S D E S I GN EYE POSITION

80 ,< 4 0 30 20 10 0 10I I LATERAL ANGLE,I I DEG

LOOKING

PARALLEL

TO Z BODY

ELEVATION

ANGLE, DEG -3 0 t I

FIGURE 10FIGURE 12

-70

NASA-SW-5140 JUNLM FLIGHT CONFIGURATION

. ."FIGURE 11

NASA-S-66 -3576 MAY 12LM DESCENT ENGINE

THRUST CHARACTERISTICS

t UPPER LIMIT

3- UNCERTAINTY

FIGURE 13 0 'TIME, SECONDS

N A S A - 5 - 6 6 - 6 4 7 0 J U N


24/39

VARIATION OF LM L A N D i N G P O S IT IO NREQUIREMENTS THAT HAVE

BEEN CONSIDERED

0 LANDING AT ANY SUITABLE POINT WITHINA SPECIFIED AREA

0 LANDING AT ANY SUITABLE POINT WITHINA SMALL AREA CONSTRAINED IN SIZEPRIMARILY BY GUIDANCE DISPERSIONS*

0 LANDING AT A PRESPECIFIED POINT (SUCHAS A SURVEYOR)

FIGURE 14 *PRESENT STRATEGY IS BASED UPON THIS REQUIREMENT

N A S A - S - 6 6 . 5 0 d d J UN

LM THREE-PHASED POWERED DESCENT

APPROACH

250 N MI0 BRAKING PHASE - ALLOWS EFFICIENT REDUCTION OF

MOST OF VELOCITY

0 FINAL APPROACH PHASE - ALLOWS ACQUISITION ANDASSESSMENT OF SITE AND CONFIRMATIONOF FLIGHT SAFETY B Y PILOT

0 LANDING PHASE-ALLOWS VERN'IER CONTROL OFPOSITION A ND VELOCITIES

LM POWERED DESCENT

TARGET SWITCHOVER

ENGINE IGNITION h = 5 0 , 0 0 0 FTh = 5 0 , o o o FT e=88o T = 2 2 8 SEC h=25; 000 FT9 = 8 6 O T = 2 8 SECT= O SECV=SSOO FT/SEC y=".2DY =o o

MAXIMUM THROTTLE h:43,000 F - . LR UPDATEe=800V - 3 3 8 5 e =71

V:5564 FT/SEC y"1.40 1 ~ 3 2 8SECVX2164 FT/SECY=-4.0

t. DOWN RANGE

FIGURE 16A

LM POWERED DESCENT ( CONT)

FICTITIOUS TARGET HIGH GATE

h = 8600 FTe=46OV-608 FT/SECT:454 SECY =-14.5O

h=16 , 000 FTLR ALTITUDE UPDATE e =4006 6 . 5 O SEC

h = 2 5 , 0 0 0 FT V=1067 FT/SECY=-4.0"-~""""""","" "-"""-~~""""-

""(2""-

Low GATE'1~558SECh=5DO FTV = 5 2FT/SEC1 = - I 7 . O 0

FIGURE 15

FIGURE 168

"..@30

"""L"";;'71~""-i- "-~25 20 15 10

. .

5 0

D O W N RANGE, N MI

N A S A - 5 - 6 6 - 6 4 7 6 J U NN A S A - S - 66 - 6425 JUN

THRUST BEHAVI OR FOR LI M I TED


25/39

BRAKING PHASE DESIGN

0 OBJECTIVES0 REDUCE VELOCITY TO ACCEPTABLE LANDING

APPROACH MAGNITUDES

0 MAINTAIN EFFICIENT USE OF PROPELLANT FUELREACH A PRESPECIFIED STATE VECTOR

AT HI-GATE POSITION

CONSTRAINTS

0 DESCENT ENGINE IS NON -THROTTLEABLEIN MAX THRUST REGION

0 MAXIMUM THRUST OF DESCENT ENGINEIS INlTlALLY=9700 LBS [ T / W % 3 )

0 FIXED THRUST UNCERTAINTIES MAY REACH 22 1/2%FIGURE 17

N A S A - 5 - 6 6 - 6 4 1 0 J U N

POWERED DESCENT IGNITION LOGIC

ACCELERATIONCOMMAND

FIGURE 18

,-ACCELERATION PREDICTED AT LITEUP

EXAGGERATED CUR-VATURE ESPECIALLY

IN THIS REGION

0

0

COMMAND ATITERATION

IGNITION TIMEIGNITE

w TIMEIGNITION TIME

THRUST BEHAVI OR FOR LI M I TEDT H R O T T L E G U I D A N C E

TRAJECTORY PROFILEo , o o o , T A R [ T i x S TARGET

RANGE FINAL APPROACH

ALTIT UDE /-HI-GATE TARGETFEET SWITCHOVER

CASE I NOMINAL T/ W AND LANDING

1 2 ,0 0 0 - \ C O M M A N D E D ( N O M IN A L )10,000 --- J. r , I II II III6000 - I \ I II4000 - I I

I I

2000 - I I0 I

I -\THRUST,LBS 8000 - I \ 'I

I I

1 I IRANGE

FIGURE 19

N A S A - 5 - 6 6 - 5 0 4 2 JUNT H R U S T F W " V l O R F OR LIMITED THROTTLE G U I D A N C Er BRAKING

TRAJECTORY A L T : T : T F mFlCTlClOUS TARGET/- HI-GATE TARGET

PROFILE FEET SWITCHOVER0

/ -COMMAND FINAL APPROACHA N D LANDING

LOW T/W P O U N D SCASE II THRUST, OMINAL5000 I

I II I ! \

,--ACTUAL

N A S A - S - 6 6 - 3 0 4 3 APR 5AV P E NA L TY D U E T O F IX ED

GUIDANCE


26/39

AV P E NA L TY D U E T O F IX EDT H R US T U N C E R T A I N T I E S

-PERCENT T/ W ERRORAV PENALTY"

80

PENALTY

FT/SECA V

GUIDANCEC O M M A N D S

0 20 40 60 80 100FIGURE 21

BIAS TIME REQUIRED FOR FICTITIOUS TARGET, SEC

NASA-5-66-6A26 JV N

0

FORPOWERED 4 0 1 A V = 6281 FPS\DESCENT

DEG20

HI-GATE 6100 FT 0PERCENT T/W

UNCERTAINTY

LAN DIN G RADAR WEIGHTING FAC TORSFOR ALTITUDE AND VELOCITY UPDATES

.8VELOCITY

WEIGHTING .4FACTORS

W V z a W Y y a

(DOWNRANGE)

400 8001200 1600

VELOCITY,FPS

ALTITUDE

WEIGHTING .4FACTOR

FIGURE 220 8 00 0 16000 24000

ALTITUDE, FT

10,000

THRUST,IDEAL CONDITION

-

LBS 50000200 300 40 0 500 600

FIGURE 23 TIME F ROM BRAKING.INITIATION, SEC

NASA -5 -66 -6513 JU N

TERRAIN PROFILE DURINGAPPROACH TO LANDING SITE

0 20' N LAT 12 50' E LONG

t 3 0 0 0t 2000

ALTITUDE, + I O 0 00

-10005EXPANDED ALTITUDE SCALEFT 25 20 15 10 5 02 O SLOPE SCALE0 PROFILE LANDINGALTITUDE, 4000 k--L---""- SITE 7FT 0-4000 t I I I I I25 20 15 I O 5 0

FIGURE 24 RANGE, N MI

NASA-S-66-6441 JUNPHASE FINAL APPROACH DESIGN


27/39

GUI DANCEC O M M A N D S

FORPOWEREDDESCENT

ID EA L CO N D ITIO N-

TYPICAL ERROR -----C O N D I T I O N SA N D T E R R A IN

FIGURE 25

-2000u8060 r : , r A V = 6 2 9 7 F P SPITCHANGLE, 40

DE G20

I 1 I , I I ,

THRUST,

LBS 5000 r F h l- "_ ,0200 300 400 500 600

TIME FROM BRA K IN G IN ITIA TIO N , SEC

N A S A 5 - 6 6 - 6 4 0 3 IU NLM PO W E R E D DESCENT GUIDANCE

MONITORING

0 P U R P O S E OF M O N I T O R I N G0 P R O V I DE A S S E S S M E N T OF T R A J E C T O R Y

0 F A I L U R E D E T E C T I O N A N D I S O L A T I O N

0 A S S U R E S A F E A B O R T

0 TWO T E C H N I Q U E S0 M O N I T O R I N G T R A J E C T O R Y B O U N D S OF P N G S A N D A G S

0 P E R I O D I C D I F F E R E N C I N G OF P N G S A N D A G S

0 ALTITUOE . A L TI T UD E R A T E M O S T S I G N I F I C A N T F O R A B O R T S A F E T Y0 A L TI T UD E R A T E D E V I A T I O N S M O S T S E N S I T I V E T O FA I L UR E D E T E C T I O N0 M S F N M E A S U R E M E N T OF A L T I T U DE R A T E S H O U L D BE S U F F I C I E N T FOR

F A I L E D S Y S T E M I S O L A T I O N

0 3 o G U I D A N C E D E V I A T I O N S WILL N O T E N D A N G E R F L IG H T P R I ORTO HI -G A TE

FIGURE 26

PHASE 1I - FINAL APPROACH DESIGN0 OBJECTIVES

PROVIDE CREW VISIBILITY OF AND ADEQUATE TIMEPROVIDE CREW OPPORTUNITY TO ASSESS FLIGHT SAFETYPROVIDE A RELATIVELY STABLE VIEWING PLATFORM

TO ASSESS LANDING AREA

0 CONSTRAINTS0 FUEL LIMITATIONS0 LM W I NDO W S I ZE0 T/W OF DESCENT AND ASCENT ENGINE AND REQUIREMENT0 TERRAIN LI G H TI NG / CO NTRAS T PROPERTIES

FOR SAFE STAGED ABORTS

VARIABLES

0 PITCH ATTITUDE

0 TRANSITION ALTITUDE

0 FLIGHT PATH ANGLELOOK ANGLE TO LANDING AREA REFERENCED TO

27 THRUST AXIS

N A S A - S - 6 6 - 6 4 0 2 JU N

FACTORS AFFECTING CHOICE OF HI-GATEALTITUDE

0 R A N G E FROM WHICH LANDING AREAC A N BE ASSESSED

0 TIME REQUIRED T O ASSESS LA NDING AREA0 FLIGHT SAFETY REQUIREM ENTS WITH

REGARDS TO TERRAIN ALTITUDEUNCERTAINTIES, LANDING RADAR

OPERATING RELIABILITY, AND ASCENT

ENGINE ABORT BOUNDARY

FIGURE 28

N A 5 A . S - 6 6 . 5 0 5 1 JUNABORT CAPABILITY BOUNDARIES

N A S A - S - 6 6 - 6 4 7 ! JUNDETERMINATION OF MINIMUM


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6000 r

ALTITUDEIFTI

5000 -

[ASSUME 4 SE C4000 - D EL AY IN ST A G IN G )

3000 -

FIGURE 290 50 100 150 200 250 30 0

DESCENT RATE IFT/SECl

N A S A - S - 6 6 . 5 0 4 1 J U N

FACTORS CONTRIBUTING TOUNCERTAINTIES IN

ALTITUDE ABOVE TERRAIN

0 GUIDANCE AND NAVIGATION UNCERTAINTIES 11500 FT ALT 101

0 LUNAR RADIUS BIAS MAGNITUDE (3200 FT ALTl0 LUNAR RADIUS RANDOM MAGNITUDE (3200 FT ALT 1.)

0 PRESENT ABILITY TO DETERMINE MARIA (FUNCTION OFAREA SLOPES (f3' 3-1 LANDINGDISPERSIONS)

0 ALLOWABLE TERRAIN VARIATI ONS WITHIN (FUN CTIO N OF+ 2 O SLOPE AND f5 % OF NOMINAL ALTITUDE LANDING

DISPERSIONS]

DETERMINATION OF MINIMUM

HI-GATE ALTITUDE WITHOUT LR UPDATING

JUALTITUDEUNCERTAINTIES,* FT

ALTITUDE BIASES, FT

ORBIT I I I I INAVIGATION S T A G E DTERRAINLUNARLUNARTERRAINPGNCS

PROFILE ABORTPROFILERADIUSRADIUS-.

MSFN

MSFN

SITE UPDATE

B LANDING

PGNCS &LANDINGSITE UPDATE

3700

3700

4500

4700

700

1000

13,700

1700

1700

9800 4300

700

80 0

3500

1800

!80 0

MINIMUM

ALTITUDE

HI-GATE

32,600

6700

7 5 0 0 I* 3U UNCERTAINTIES ARE ROOT-SUM. SQUARED

FIGURE 31

N A S A . 5 - 6 6 . 6 1 3 3 JUN

25 0 r NO M IN A LH I- GA TE ~ O O G T-AV PENALTY

FOR HI-GATEALTITUDE

VARIATION-TYPICAL

FLIGHT PATH

ANGLE = 15"

200

A VFPS

150

100

50

0 2000 4000 6000 8000 10.000HI-GATE ALTITUDE, FT

FIGURE M FIGURE 32

NASA 5 6 6 6 4 2 0 IUNAV PENALTY FOR LOOK A N G L E A N D F L I G H T

N A S A - S - 6 6 . 6 5 7 9 JU NOVERHEAD PROFILE OF FOOTPRINT CAPABILITY

FROM 8000 ALTITUDE


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PATH ANGLE (HI -GATE 9000FT)

1100

FT / 5 E C

- 100FIGURE 33

LOOKA N G L E

T OL A N D I N GS I T E , D E G-c - 3 0 W I N D O WL O W E R- 2 5 LIMITI I 1 I 15 10 15 2 0 2 5

FLIG H T PA TH A N G L E , D E G

N A S A - S . 6 6 - 6 I 9 5 JUNCOM PARI SON OF DESIGN TRAJECTORY

A N D FUEL OPTIMUM

20 0 r 8 6 0 0 FT,ALT 7VERTICALVELOCITY 10 0

FT/SEC 50I I I 1

0 20 40 6 0 80 10 01500 r TIME, SEC - N O M I N A LE S IGN

-- FU EL O PTI MU M

!

FROM 8000 FT ALTITUDED O W N R A NG E

3 0 X l O I , r S l D E yAfpOW

3000

ENCE

FIGURE 35

A V ,F P S

FT REFER-

L O W E R W I N D O W L IM IT

PO SITIO N O V ER SU RFA CE A TTIME OF R E D E S I G N A T I O N

N A S A - S - 6 6 - 3 2 9 1 APR 16VARIATION OF FOOTPRINTCAPABILITY WITH ALTITUDE

20x103 D O W N RANGEI F O O T P R I N T S F O R

10 1 A- AV,=lOO

A A z = 3 0 / \:;:% AT A L T z 3 0 0 0 FTP O S I T I O N O V E R S U R F A C ELP O S I T I O N O V E R S U R F A C EA T A L T = 5 0 0 0 FT-30 A T A L T = 8 0 0 0 FTFIGURE 34 T IME ,S E C

FIGURE 36


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lunar landing symposium

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