E -2476

MAN-MACHINE DESIGN FORTHE APOLLONAVIGATION, GUIDANCE, ANDCONTROL SYSTEM - REVISITED:

SUBTITLE

APOLLO, ATRANSITION INTHE ARTOFPILOTING AVEHICLE

byJ.L. Nevins

JANUARY 1970

C A M B R I D G E . M A S S A C H U S E T T S , 0 2 1 3 9

ACKNOWLJEDGMENT

This report was prepared under DSR Project 55-238’70, sponsored by the

Manned Spacecraft Center of the National Aeronautics and Space Administration

through Contract NAS g-4065, with the Instrumentation Laboratory, Massachusetts

Institute of Technology, Cambridge, Massachusetts.

The author extends his thanks to the following members of the Instrumenta-

tion Laboratory.

For technical criticism - Mr. Ivan Johnson and Mr. John Scanlon

For editing - Mr. Jack Reed

For publishing - Mr. Robert Weatherbee and associates, particularly

lVIr. Wallis Bean for editorial support.

The publication of this report does not constitute approval by the National

Aeronautics and Space Administration of the findings or the conclusions contained

therein. It is published only for the exchange and stimulation of ideas.

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Man-Machine Design for the Apol lo Navigat ion,Guidance, and Control System-Revisited:

Sub-tit leApol lo, A Transit ion in the Art of Pi lot ing a Vehicle

J. L. Nevins

I INTRODUCTION

Apollo can be considered a transit ion in the artof pi lot ing avehicle, where the principal dimensionsa r e (a) F l i g h t O p e r a t i o n s , (b) t h e F l i g h t C r e w ’ sRole, and (c) the man-machine communications, asi l lustrated by Fig. 1. The transitional aspects arethe level and the nature of integration of the aircrewsand ground control lers for f l ight operations. F o rthe crews, the aspects are (1) the spectrum, or rangeof levels , o f the general tasks and the necessity forcertain tasks, (2) the nature and the requirementso f t h e s u p e r v i s o r y r o l e . For the man-machinecommunications, the signi f icant i tems are the levelsand the nature of interaction of the crew with theirequipment, from direct actuation of effecters to af irst level of functional communications.

Consider, for example, the primary guidance,navigation, and control system designed for theApollovehicles. The system was designed to providethe crew with a complete onboard f l ight-managementsystem that would enable them to navigate and guidetheir spacecraft without ground assistance. As such,Apollo is the first manned U. S. spacecraft to containenough sensors and data processing capabil ity to dothe job.

A . F L I G H T O P E R A T I O N S

Apol lo is the culminat ion of development both inground and airborne systems. Ground systems (Fig.1) have progressed from a few people giving minimalassistance to airplane crews (beginning in the 30’s)to relatively advanced systems for military “com-mand and control”, such as the continental air defensenetwork for North America. Systems that essential lygive only directional data, however, are signif icantlydif ferent from the systems developed for supportingmanned spacecraft: the systems supporting Apollomonitor al l spacecraft systems, and, in ef fect, theground control lers feel as though they are insidethe vehicle” and are giving direct support to f l ightoperations. This ground support ranges al l the wayfrom sequencing the proper charging of batteries,to trajectory control in scheduling the small thrustingef fects f o r w a s t e - w a t e r d u m p s . T h e flight-demonstrated capabil i ty of the ground’s monitoringability, p l u s t h e n e a r - p e r f e c t r e l i a b i l i t y o f t h e

*Modif ied by the fact that the data are “old,” bothbecause of transmission delay and because of systemdelays in the ground communication system andassociated processors.

onboard equipment, give the aircrews the czzfidenceto rest without keeping one man on watch.

Spaceborne systems for manned spacecraft haveprogressed from the minimal onboard capabil ity ofMercury, through Gemini with onboard navigationand guidance capabil i ty for rendezvous, to Apollowith ful l onboard capabil ity for performing the fulllunar-landing mission.

To achieve the desired mission rel iabi l i ty goals,the f l ight-management system, instead of beingprimarily an onboard operation, is actually a highlyintegrated system of airborne and ground-basedequipment . The nature o f th is integrated team is af inely structured multi level monitoring and decisionprocess. In its simplest mode, the aircrew monitorsthe datafor detecting errors that require immediateaction, while the ground control lers are responsiblefor detecting the gradual-degradation-type fai lureo f the onboard sensors. The latter fai lures can onlyb e d e t e c t e d o n t h e g r o u n d b y m o n i t o r i n g a n dcomparing the long-term trends of the databoth fromthe airborne and ground-tracking systems.

In addition, since Apollo was man’s f irst ventureinto deep space, maximum support was organizedon the ground to help with any contingency. Thissupport included not only the people manning theconsoles in Houston, but hundreds of people at thevarious contractor faci l i t ies around the country(North American Rockwell in Downey, Cali fornia;Grumman Aircraft and Engineering Corporation, onL o n g I s l a n d , N . Y . ; t h e M . I . T . I n s t r u m e n t a t i o nLaboratory, etc. ) , a l l t ied together by voice- anddata-communication l inks. Marshall ing this kind ofsupport in depth would be impractical i f we wereflying multiple missions simultaneously; e.g., alunar-landing mission, an earth-orbit-equatoriallong-duration mission, and an earth-orbit-polarmission, al l manned. There fore , ground- supportsystems for future manned missions can be expectedeither to become more automatic or e lse airbornesystems wil l become more autonomous. The l a t t e rtechnique is necessary at distances where trans-mission delays are minutes long. Under these-**Another example is leaving the LM vehic le unat-tended whi le both crewmen are explor ing the moon.At f irst reading, this would to appear break one ofthe old explorers ’ prime ground rules; namely, (a)never leave a vehic le unattended, or (b) neve r l e t aman explore a lone. The ground in this case real lyacts as a third crew member to monitor the LMwhi le the other crew members explore the moon.

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condit ions, ground systems for space operat ionscould only support airborne operations in the samew a y t h a t t h e p r e s e n t g r o u n d s y s t e m s s u p p o r tairplanes in f l ight.

B . F L I G H T C R E W ’ S R G L E

1 . G e n e r a l T a s k s . - M a n ’ s r o l e i n s p a c e c r a f tguidance and navigation ranges from supervisingautomatic systems to performing specif ic sensingand control functions. The crew funct ions can becategorized as fol lows:

a. Monitoring of , and decision making associatedwith, t h e n a v i g a t i o n a n d g u i d a n c e p r o c e s s ,including the ef fects of navigation sensor data(target-tracking data, both visual and radar) onstate-vector updates; comparing onboard datawith ground-tracking data and backup charts

b . S e q u e n c i n g a n d i n i t i a l i z a t i o n o f p r i m a r yguidance, navigat ion, sensing systems, as wellas propulsion and timing systems

c. Init ial izing and sequencing of backup systems

d . P e r f o r m i n g t h e p a t t e r n - r e c o g n i t i o n t a s k sassociated w i t h (1) c o m m a n d - m o d u l e o p t i c a ltracking of the lunar module during rendezvous;(2) star acquisit ion, identi f ication, and geo-metrical al ignment to v isual horizon (earth ormoon) for cislunar navigation.

As technology improves in capabil ity and rel ia-bility, m a n y o f t h e s e t a s k s w i l l b e r e p l a c e d b yautomatic systems. In Apol lo, however, in manyrespects the most complex vehicle every pi loted,success depended upon a design that thoroughlyintegrated man and machine, a design concept thatu t i l i z e d m a n t o a c h i e v e s y s t e m f l e x i b i l i t y a n drel iabi l i ty not otherwise possible given presenttechnology.

2. Supervisory tasks.- The rel iabi l i ty of Apol loequipment demonstrated industry’s abil ity to producesystems that meet specif ied goals. Nevertheless,the l imited rel iabi l i ty of basic components, togetherwith the constraints on weight, volume, and power,produced system designs where single fai lures cancause large funct ional incapacitat ion of the af fectedsystems. To guarantee funct ional capabi l i ty , re-d u n d a n t s y s t e m s a r e n e c e s s a r y . M a n ’ s m o s timportant role, therefore, especial ly during dynamiccondit ions, is to monitor both the primary systemand its required backup systems. F o r c r i t i c a lfunctions, this requires that the crew give continual,t ime-shared attention to several levels of backupsystems in order that their status be known shouldtheir use become necessary. Moreover, smooth andrapid transit ion to backup modes requires the crewfunct ional involvement in the operat ion of the tota lsystem.

Awareness of , and involvement in, the operat ionof many levels of redundant systems, operating inparal le l , places a most severe burden on the crew.With increased rel iabil i ty and smaller size of basiccomponents of the future, i t wi l l be possible toprovide enough redundant sensors, electronics,processors, and highly rel iable switching logic tonot only detect malfunctions, but to automatical ly

switch to redundant modes-that is, a system thatdegrades gracefully rather than instantly (Ref. 1).Such a system would operate most o f the t ime in afully automatic mode. For the next five to ten years,however, i t is unl ikely that man’s present uniquef lexibi l i ty a n d d e c i s i o n c a p a b i l i t y c a n b e f u l l yreplaced; during this t ime, systems wil l have to beconf igured to al low man’s continued involvement atlevels other than required str ict ly for supervisingor monitoring. The depth of this involvement,however , s h o u l d b e m u c h l e s s t h a n i n A p o l l o .Consequently , the burden on the crew wi l l cont inuet o d e c r e a s e a s i t s r o l e b e c o m e s m o r e p u r e l yadministrative, or supervisory.

C . M A N - M A C H I N E C O M M U N I C A T I O N

Communicat ion between crew and the airborneequipment comprises everything from direct tasksequencing, caused by direct actuation of effectersvia hand control lers, to a functional level of com-munications implemented by a higher -level computerlanguage (Fig. 1 and ‘71. This computer languageal lows the crew to control groups of tasks insteadof individual tasks. These task groups can be assmall as an automatic spacecraft-att itude rotationand as large as required to integrate al l the jobsnecessary f o r p e r f o r m i n g t h e p o w e r e d - d e s c e n tportion of the lunar landing or entry into the earth’satmosphere. Although this language is a majorcontribution to the art of pi lot ing a vehicle, i t iscrude by the standards of newer technology that allowv a r i a b l e f o r m a t , graphical input/output displaysystems. In the remainder of this paper, I wil l f irstdiscuss the Apollo mar-machine interface from thebroad perspective of the primary guidance, nav-igation, and control system. I will then discuss thesal ient features of the Apol lo f l ight-managementsystem, including mar-machine communications, asapplied to lunar-landing powered descent; and, final-ly, the mar-machine control interact ion for thecommand module rotational att i tude control modes.

11 BROAD VIEW OF APOLLO MAN-MACHINEI N T E G R A T I O N

The Apol lo Guidance, Navigation and ControlSystem is described in Ref. 2 through 13. T h eprincipal aspects of the man-machine interaction aredescribed in Ref. 14, 15, and 16. A few words arenecessary, however, to provide a context for theremainder of this paper.

To control the spacecraft throughout the basiclunar- landing m i s s i o n ( F i g . 2 ) e n t a i l s f i f t e e n *distinct operational phases for the onboard Guidance,Navigation, and Control System. To accomplishthese funct ions in the l ight o f the pert inent groundrules (Table 1) requires a highly integrated systemwhose primary inputs are shown functionally by Fig.3. Onboard navigation data come from three sensors.Two are on the command module, namely optics anda r a n g e - o n l y c a p a b i l i t y t h r o u g h t h e V H F c o m -munications link. On the lunar module, the navigationsensor is a radar system.

“Note: Orbital navigation (earth and moon) ( i tems2 and 6 on Fig. 2) are not normally used, andmidcourse navigation (items 4 and 13, Fig. 21 areonly uti l ized in the “ loss-of-communication” case .

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Table 1 Apollo Design Ground Rules (circa 1961)

a. The system should be capable of completingthe mission with no aid from the ground; i.e.,self-contained.

b. The system will effectively employ humanparticipation whenever it can simplify orimprove the operation over that obtained byautomatic sequences of the required functions.

c. The system shall provide adequate pilotdisplays and methods for pilot guidance systemcontrol.

d. The system shall be designed such that onecrew member can perform all functionsrequired to accomplish a safe return to earthfrom any point in the mission.

The optics system, on the command module, isdesigned to provide navigation data in local orbits(earth or moon) by single-line-of-sight observations(via a one-power optical instrument called the scan-ning telescope-SCT) from the inertial platform toknown or unknown landmarks. In cislunar spacenavigation, data are obtained by two-line-of-sightobservations (via a 28-power optical instrumentcalled the space sextant-SXT) between stars andearth or moon horizon. During rendezvous, the opticssystem and the range-only capability through theVHF communications radio link provide the naviga-tion data on the command module; on the lunarmodule, the navigation data come from therendezvous radar. State vectors derived from groundtracking can be sent to either spacecraft during anycoasting-flight phase, via telemetery. Guidance ofthe Apollo spacecraft is inertial; i.e., applied forcei s sensed by accelerometers mounted on agyroscopically stabilized platform and processed bya computer that generates steering and engine-cutoffcommands, as shown functionally in Fig. 3. Thelunar module G&N system also uses radar andastronaut-visual inputs during the final approach tolanding, and therefore the LM may be said to useradar-visual inertial guidance (Fig. 3). Thecomponents for the command module (CM) primaryguidance, navigation, and control system (PGNCS)with their respective locations are illustrated by Fig.4. In Fig. 5, the lunar module (L,M) is similarlydetailed. The onboard computers, which areidentical, are identified as the command modulecomputer (CMC) for the command module and thelunar module guidance computer (LGC) for the lunarmodule. These processors are the primary onboardsequence controllers as well as the clock or basictime and frequency reference for the spacecraft.Figure 6 shows the interrelationship of the LGC tothe various sensors, the spacecraft reaction- controlmotors, and the spacecraft propulsion system forthe LM digital autopilot functions. The computersperform the following: (1) monitor the sensor data(optics, accelerometers, IMU gimbal angles on thecommand module, rendezvous radar, landing radar,accelerometers, and IMU gimbal angles on the lunarmodule); (2) determine thrust times and vectors,vehicle-trajectory parameters, and optics or ren-dezvous radar-target lines of sight; (3) maintainattitude control; and (4) guide the vehicles duringthrusting maneuvers.

The computer and crew primarily interface atthe display/keyboard (DSKY) (Fig. 7a), whichconsists of electroluminescent digital displays anda numeric keyboard. Data are displayed in threefive-digit registers. The displayed data can be eitherdecimal or octal, Associated with each register isa sign bit for the display of decimal data. In addition,memory locations can be addressed directly, but thisis intended primarily for ground checkout. (Althoughthere is ng attempt to restrict crew access to thecomputer, the crews are trained to use primarilythe technique designed for flight operations.)

For flight operations, crew-computer com-munications are primarily structured into two levels(Ref. 17). The first level identifies GN&C operationalfunctions (not usually as large as a mission phase-Fig. 21, e.g., (1) targeting computations for one ofthe four rendezvous subphases, (2) the sequencinga.ssociated with a particular propulsion system fora. trajectory maneuver. This highest-level functionalidentifier is a two-digit decimal code called aprogram (P) identifier, where the most significantdigit is related to mission phase. Figure 7b il-lustrates the organization. For example, zero seriesidentify the functions associated with prelaunchcheckout, the ten series identify the boost-monitoringprograms, and the sixty series identify programsassociated with LM lunar landing and CM entry.The unit programs define the functional programswithin a particular series. Figure 8 gives thecomplete list of programs for a typical lunar-landingmission.

The second level of communication consists oftwo, two-decimal, digit identifiers called verbs andnouns. The verb identifier defines action and theaction to be performed; the noun identifier definesthe object of the action and identifies the data beingdisplayed or loaded. (Figure 7b lists some typicalverbs and nouns.) For example, Verb 16 instructsthe computer to continuously monitor a function,display the data in decimal form, and update thedisplay every two seconds. If we combine Verb 16with Noun 36, we will instruct a display of thecomputer clock, expressed in ground elapsed time(GET), with hours in the first dataregister, minutesin the second data register, and seconds to hundredthsof a second in the third data register.

Communication with a computer is alwaysbimodal, i.e., the man talking to the computer, andt.he computer talking to the man. The latter modeis mechanized by allowing the computer to flash theverb-noun displays (flash rate l/2 second on, l/ 2second off). Therefore, if the computer wants theman to review data for acceptance or rejection orto load data, it will flash the appropriate verb-nouncombination.*This philosophy gives great flexibility, but has theassociate hazard of a possible wipe-out or incorrectalteration of the data in erasable memory (statevector, etc.) if the computer is not sequencedproperly. For example, on the transearth leg ofthe Apollo 8 mission, a very tired crew memberinadvertently called for a prelaunch function(Program 01-see Fig. 7) to be performed when heactually intended to load Star Number 01. To guardagainst these contingencies, the crew carries dataand procedures to reinitialize the erasable memory.

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In addition, there are activity lights for both thecomputer and the telemetry uplink, as well as cautionand warning annunciators for both the computer andthe rest of the inertial system.

Figure 9 shows the general nature of the crew-computer interface for per forming operat iona lsequences. This design permits the crewto exercisecomplete sequence control but also results in thecrew’s being an integra l part o f any computersequencing. Figure 15 detai ls the computersequencing for the lunar-landing powered-descentphase. Additional description of the operationalman-computer interface will be given in the lunar-landing section.

HI LUNAR LANDING PHASE

To illustrate the salient features of the presentApol lo f l ight-management system, I will nowdescribe in some detai l the powered-descentfunctions for a lunar landing. Figure 10 gives theabbreviated lunar-orbit time line for a typical lunar -landing mission. The time period covered (elapsedtime of approximately 55 hours) is from the secondlunar-orbit-insertion trajectory maneuver throughthe transearth-injection trajectory maneuver. Forthis discussion, the principal interest is the lastl-1 / 2 hours before touchdown (from separation ofthe two vehic les to the landing) -nominal groundelapsed t ime (GET) of 1Ol:OO hours , as shown inFig. 11 and 10. Figures 12, 13, and 14 show theprincipal operational phases and events associatedwith the lunar-module powered descent; Figure 15is a profile view of the powered descent, includinga listing of the computer-program-sequenceddisplays and the pertinent nonsequence displays thatcan be called by the crew.

The early phase of lunar-orbit tasks starts aftercompletion of the lunar parking-orbit circularizationmaneuver and ends after successful completion ofthe descent-orbit maneuver and the associated post-burn (or maneuver) checks. During this period thelunar module is partially activated and the command-module pilot practices optical tracking of the landingsite by tracking lunar landmarks. The crew thenrest for nine hours or so. After the crew rest period,the two lunar-module crewmen, with detailed supportf r o m t h e M i s s i o n C o n t r o l C e n t e r p e r s o n n e l ,completely act ivate and check out a l l the lunarmodule systems. Included in these checks is a dumpof the erasable memory (2,048 sixteen-bit words)in order to allow a detailed check by the ground ofthe initial load in the memory. During this checkout,al l the backup systems are activated and checkedwith the primary systems and cross-checked withthe ground. For GN&C, there is a complete backupsystem (abort guidance system- AGS) that is capableof guiding an abort f rom any place during thepowered-descent trajectory. The AGS is also capableof guiding the rendezvous maneuvers.

After the lunar-module checkout, the command-module pi lot (CMP) tracks the landing s i te (or aprominent landmark near the landing site) with his,optics system. These data are sent to the groundfor processing by the Miss ion Control Centercomputers in order to reduce the relative uncertainty

between the landing- site location and the. CSM orbit.Next the vehicles undock and then separate by thecommand module’s doing a small (2: 2.5ft/ set) radialmaneuver to put it on an equiperiod orbit with thelunar module. This maneuver results in a maximumseparation of a couple of nautical miles, which occursone-hal f orbit later at descent-orbit insert ion.During the next night pass, the inertial platformsof both vehic les are carefu l ly a l igned, us ing twostars. A check to see if the proper stars were usedis made by an auto-optics routine for the commandmodule (auto-spacecraft routine for the lunar modulebecause its optics are fixed); this routine points theoptics or vehicle to a third star. The gyro torquing-to eliminate the gyro drift since the last alignment-indicates whether the gyro-drift performance is goodenough to continue the landing. Additional checksare made on the bias terms for the accelerometersand to see if the rendezvous radar on the LM andthe VHF range channel on the CM are workingproperly. While these checks are going on, the groundcontinually monitors the other spacecraft systems(except for the time period when the spacecraft arebehind the moon) to see if they are operating andsequencing properly.

The LM crew then se lect the ground-targetedpre-thrust ing program (P30). After sequencingthrough P30, they then select the descent-propulsion-system computer program (P40); maneuver to thespacecraft att itude required for the Hohmanndescent-orbit maneuver; andverify that the computerdisplays for velocity components, time of ignition,and spacecraft attitude all agree with the ground-computed data voiced-linked from the ground. Theignit ion att i tude is ver i f ied by looking out theoverhead window at the star selected by, and voicedfrom, the ground.

The descent-orbit - insert ion (DOI) maneuver isperformed about 194 longitude degrees before thetargeted landing site. The result of this maneuveris to place the lunar module in a 60-nautical-mileby 50,000-foot Hohmann transfer orbit, with periluneoccuring about 14 longitude degrees before thetargeted landing site. At perilune, powered descentwill be initiated.

After the maneuver is performed, the results ofboth the primary and the backup system and theindicated rendezvous-radar range rate between thetwo vehicles are evaluated to verify that the maneuverwas successful. During the maneuver, if the descentengine stayed on three seconds longer than it shouldhave, an unsafe perilune would result.

The principal decisions possible during this finalphase are as fo l lows: (a ) GO/NO GO for landing;(b) if an abort is required, should the abort be madeon the descent engine, or should the vehicle be staged(separate the descent stage from the ascent stage)and the abort be made on the ascent engine; and (c)whether to switch over from the primary guidancesystem to the backup system in case of a failure inthe primary system. For an acceptable conditionfor landing, there must be close agreement betweenthe onboard-computed values a n d t h e v a l u e scomputed on the ground and voiced up to the crewfor timeof engine ignition and cross-range location

o f t h e l a n d i n g s i t e t o t h e L M o r b i t a l p l a n e . I naddit ion, the LM inertial-platform alignment mustbe checked, the LM must achieve the proper pitchattitude for the trajectory maneuver, and the airbornecomputer must be functioning properly.

As indicated earl ier, the ground is responsiblefor detecting insidious slow-drift sensor malfunc-tions that would require switchover from the primarysystem to the backup system. To do this, the groundhas a special powered-f l ight processor, which usesa Kalman f i l ter ing technique to generate a statevector, by processing the doppler-range data andt h e g e o m e t r i c a l - t r a n g u l a t i o n d a t a f r o m s e v e r a ltracking stat ions. This statevector is then comparedwith the telemetered onboard-computed state vector.In addit ion, g r o u n d d i s p l a y s s h o w t h e v e l o c i t yr e s i d u a l s b e t w e e n onboard p r i m a r y a n d b a c k u pguidance systems, onboard primary guidance systemand ground-tracking-derived range-rate residuals,and onboard backup and ground tracking-derivedrange-rate residuals. For these comparisons to bemade requires that there be continuous high-bit-ratedata and voice communicat ions between the groundand spacecraft.

As indicated in Fig. 12, 13, and 15, there arethree main phases to powered descent.

a. Braking phaseb. Approach or visibility phasec. Landing or vert ical descent phase.

The braking phase (Computer Program P63) startsnear per i lune (- 50,000 ft) o f the Hohmann descent-transfer orbit, about 260 nautical miles away fromthe targeted landing site. The braking phase lastsapproximately 8 min 30 set and ends when the targetconditions (called HI-GATE) for thevisibil ity phasehave been met. During this braking phase, the LAIis in a retrograde attitude and is using the full thrustof the descent-propulsion system to s low down thevehicle. About three-quarters (nominal t ime about6:30 min - 24,000-ft altitude) of the way through theburn, the guidance law throttles the engine down toapproximately 57 percent. Before thrott le-downtime, at an alt itude of approximately 40,000 ft , thespacecraft is manually rotated to a windows-up (F ig.121 att i tude ( i f required) in order that the landingradar can lock on the lunar surface. At about 8.22,the vehic le begins an autopitch maneuver to enablethe crew to view the landing site through the forwardwindow during the next phase-the v is ibi l i ty phase.

Before this, the guidance law has been slowlypi tching the vehic le to enable the engine thrust tomaintain the proper alt i tude prof i le as the vehicleslows down. At approximately 8:30 min (computerp r o g r a m P64), t h e n e x t p h a s e b e g i n s - t h e final-approach, or landing- site-visibi l i ty phase. Thisvisibi l i ty-phase program starts when the HI-GATEtarget condit ions have been met (alt i tude “7,200 ft ,inert ial velocity w 516 ft/sec, fuel remaining ap-proximately 20-percent) and ends about 120 secondslater, when the LO-G ATE target conditions (altitudeN 150 ft, vertical-descent velocity 3 ft/sec, ap-proximately lo-percent fuel remaining) have beenmet. During this visibi l i ty phase, the computerdisplays an elevation angle (via DSKY 1 that indicatesto the crew where the computer p lans to land thespacecraft.

The crewman surveys the lunar terrain and, byuse of a s imple ret ic le on the window, notes wherethe spacecraft is going to land, I f the crewman isnot sat isf ied with the pretargeted landing-site ter-rain features, he can designate another landing site,or he can f ly the spacecraft to another landing sitein either a semi- auto mode (compu.ter program P66)or a fully manual mode (computer program P67).These modes wi l l be descr ibed later . At the end ofthe vis ibi l i ty phase, i f the crewman does not e lectP66 or P67, the autolanding program (Pfi5) will beautomatical ly entered.

In P65, the vehicle is pitched to the vert icalposit ion, the translat ional velocit ies are nulled, andthe vert ical-descent rate is set to 3 ft/sec. Justbefore touchdown, a lunar contact l ight is act ivatedb,y 5-foot- long probes attached to the landing pads.When the lunar contact l ight comes on, the crewmanually shut down the engine and the vehicle fallsthe remaining distance to the lunar surface (nominalimpact velocity < 3ftl secl.

To start the powered-descent sequence, the crewcall up the braking-phase program (P63) about 40minutes before the time to turn on the braking engine(Fig. 14). This is accomplished by a Verb 37 (changeprogram request-Fig. 7b) fo l lowed by an enter (Eland a 63E as shown in Fig. 15. The f irst display inthis program is the computation result for the ignitionalgorithm. T h e f l a s h i n g V e r b 0 6 ,and N o u n 6 1displays the fol lowing in Registers 1, 2, and 3:

Rl - T ime to go(TTG), the computed t ime fromengine on to the t ime the LM wi l l reachthe HI-GATE target

R2 - T ime fromignition(TFl), the t ime to whenthe engine wi l l be automatical ly turned on

R3 - Cross Range, the distance out of plane fromthe present posi t ionto the targeted landings i te .

-Xs mdicated b e f o r e , the computer wil l f lashdrsplays r e q u i r i n g c r e w a c t i o n - m t h i s c a s e , t h eacceptance or reject ion of data. These d isplayeddata are checked against the data computed on theground and voiced to the crew. After checking, thecrew accept the data by te l l ing the computer top r o c e e d t o t h e n e x t s t e p i n t h e s e q u e n c e - b ydepressing the proceed key (PRO) on the keyboard(T:ig. 7al.

T h e n e x t s t e p m the sequence requires thecomputer to ask the crew a question. To accomplishthis the computer f lashes a Verb 50 (please perform)Noun 25 (checklist i tem), where the particular taskis identi f ied by a number displayed in register oneCR1 ). For this step, the Rl code is 00014, whichmeans, “ d o y o u w a n t t o f i n e a l i g n t h e i n e r t i a lplatform?” Since he has already previously al ignedthe platform, the crewmember bypasses this requestb y d e p r e s s i n g t h e e n t e r k e y (El. I f he had notperformed the al ignment, he would depress theproceed key (PRO). Depressing the PRO key startsthe f ine-al ign computer routine (R57). The en t e r(El causes the computer to proceed to the next task,which is an automaneuver of the spacecraft to theattitude required for starting the braking-trajectory

maneuver. This task is designateti as theautomaneuver subroutine (R60). The first displayin R60 is a flashing Verb 50 (please perform) Noun18 (the inertial referenced angles that will be seenby the crew on the three-axis-ball-attitude displayafter the computer has rotated the spacecraft to thedesired attitude-Fig. 5). Proceeding on this displayactivates the automaneuver and causes the displayto change to a static Verb 06 Noun 18. When thecomputer completes the maneuver, the display willreturn to a flashing V50N18 display. The crew thencheck to see that the DSKY, the ball-attitude display,and the ground data all agree. To exit this routine,the enter key is depressed. Note that the crew couldhave done the maneuver manually by bypassing theroutinewithanenter (E) thefirst time Verb 50 Noun18 was displayed.

The next display is a static Verb 06 (displayeddecimal) Noun 62:

Rl -

R2 -

R3 -

Present LM inertial velocity (VI)

Time from DPS engine ignition (TFI)

The measured change in velocity (AVM),which is displayed until the engine-startingsequence begins.

At TFI equals minus 35 set, the entire DSKYdisplay is blanked for five seconds to indicate tothe crew that the computer is starting the jobsassociated with a thrusting maneuver. At TFI equalsminus 30 set, Verb 06 Noun 62 returns, and thecomputer starts monitoring the accelerometers, andany residual noise is indicated by small changes inthe V06N62 display. A “run-away,” or failed, ac-celerometer would be indicated by large changes inthis display. At TFI equals minus 7 set, the computerperforms ullage by turning on the +X reaction-controlmotors. At TFI equals minus 5 set, the computerrequests (flashing Verb 99 Noun 62) a final checkfrom the crew before turning on the DPS engine. Ifthe crewmember is satisfied that all spacecraftsystems are GO, he depresses PRO, and the displayreturns to a static V06N62; at TFI equal 0 set, theDPS engineignites. In addition, the abort programs(P70 for DPS and P71 for APS) are now availableshould the crew depress the abort or abort-stagebuttons. If something were wrong at TFI=O, thecrew would terminate the entire braking sequenceby the terminate verb (Verb 34E), exit P63, and goto the idling program (POO). By not depressing thePRO key, the crew can slip starting engine ignitionfor approximately 5 set beyond the nominal starttime. Beyond the 5-set slip, mission rules requiredelaying powered-descent insertion (PDI) for oneorbit or aborting the landing, depending on theparticular spacecraft problem that caused the crewto delay the engine-start sequence. Depressing thePRO key, once TFI has passed zero, will im-mediately start the engine.

At engine start (TFI=O), the DSKY display changesto a static Verb 06 Noun 63, where Rl is the sameas Noun 62, namely inertialvelocity; R2 is the altituderate, or verticalvelocity; and R3 is the altitude abovethenominal-landing-site lunar radius. Because the

engine-gimbaling drive motors are very slowmoving, the descent engine comes on at lo-percentof full thrust for 26 seconds. During this time, thedigital autopilot positions the engine gimbal such thatthe thrust axis passes through the vehicle center ofgravity. At TFI = +26 set, the computer commandsloo-percent thrust from the DPS engine.

During the first four minutes” of the brakingphase, the crew and the ground monitor all thespacecraft systems for proper operation, and theground compares the various guidance system’s stateve-:tors to ascertain if there are any drifts in theonboard sensors. Figures 16a and b show typicaldata formats for the ground consoles for monitoringthe onboard guidance, navigation, and controlsystems. Figures 17a, b, and c show a simplifiedform of checklist that the airborne crew use tomonitor their systems.

During this period, the crew can have their “front”windows facing down or un. For Apollo 11, the crewhad their winlows facing the moon (down) and usedthe reticle (called the landing-point designator-LPD) (Fig. 5 and 18) to monitor landmarks as theypassed under the LM. At about two minutes intothe burn (Fig. 17a1, the crew also checked the altitudeprofile of their trajectory by observing how fast alandmark swept across the window reticle.

At about 40,000 feet (time = 3.0 min),** if thewindows are facing the moon, the crew would yawtheir spacecraft right 174 deg in order to orient thewindows approximately up a& enable landing-radarlockon. The crew would not vaw the full 180 deg,because of the following: downjto 30,000 ft, the crewhas control of spacecraft rotations about the thrustaxis (yaw about spacecraft X-axis, called X-axisoverride); at 30,000 ft, however, the computer locksout this capability in order that it can accuratelypoint the landing-radar antenna. By leaving in asmall error (6 deg) for the computer to null, thecrew determines that automatic X-axis control at30,000 ft does in fact occur.

Just before four mmutes, all the ground-monitor

R ersonnelG O ”

check their displays and give a positiveto the flight director, who in turn gives it to

the Capsule Communicator (CAPCOM), who relaysit to the crew (Fig. 17a). This positive “GO” isrepeated once more at about 3,000-ft altitude, nineminutes into the burn (Fig. 17b).

The next event concerns the proper operation ofthe landing radar. In addition to the operational

“The crew can update the position of the targetedlanding site with the latest data from ground trackingby use of loadable Noun (N69). He can do this anytimeduring the landing (down to P64) but normally doesit within the first three minutes after engine ignition.A V25N69E calls Noun 69 which can then be loadedwith the AX, AY, AZ components of the desiredchange in the landing site location.

**Note: Time is referenced to the time the brakingphase started.

checks by the computer and the ground,” the datam u s t b e c o m p a r e d w i t h t h e c o m p u t e r i n e r t i a l -alt i tude data before the crew wi l l a l low data to beinserted into the computer to modify the inert ial-alt i tude data. I f the di f ference is too large, theresulting trajectory could impact the moon (Fig. 19).T O perform this check onboard, the crew cal l Verb16 (monitor decimal) Noun 68, where Rl is slantrange to landing site; R2 is TG, t ime remaining inthis phase**; R 3 i s AH, t h e d i f f e r e n c e b e t w e e nalt i tude computed from the landing radar data andthe inert ia l a l t i tude. I f the AH data are within theprescribed l imits (Fig. 19) for ten seconds, the crewenables the updat ing of the inert ia l a l t i tude withlanding-radar data by performing Verb 57E. Afterenabling the landing-radar updating, the f l ight crewand the ground monitor (Fig. 16) to see that the &Iconverges as the landing- radar data force the inertiala l t i tude to agree with the true a l t i tude. Oncesatisfied, the crew can go back to the Noun 63 display,from the Noun 68 display, by depressing the KEY-RELEASE key, which allows the computer to displaythenominal-sequence display, in this case Noun 63.

The next event is the throt t le down t ime, whichis predicted by ground processor and checked bythe airborne crew. After that, the crew monitorst h e pitchup a n d t h e a u t o m a t i c s e l e c t i o n o f t h evisibility-phase program (P64) when the HI -GATEconditions have been met.

When P64 is cal led, the DSKY display changesfrom a static Noun 63 to a f lashing Verb 06 Noun64:

Rl - Time to go (or remaining) in this phase (TG)plus the landing-point-designator e levat ionangle (LPD)

R2 - Alt itude rate

R3 - Altitude

As described earlier, during this phase the vehicleis pitched up enough that the targeted landing sitecan beviewed out the front windows (Fig. 18). Also,the targeted landing-si te e levat ion, indicated by theelevat ion angle displayed in the r ight two digits ofRI, can be viewed by the crew by using the windowreticle (Fig. 18). I f the crewmember does not l ikethe indicated landing s i te , he can redesignate thelanding site by the fol lowing procedure. First, heenables redesignation by depressing the PRO key,which changes the display from a f lashing Verb 06Noun 64 to a static Verb 06 Noun 64, indicatingcomputer activation of this capabil i ty. The crewmancan then use his three- axis- rotational hand control-ler (RHC) to redesignate by moving the control lerout of detent in the desired direction and then letting

*It should also be noted that the landing radar data(alt i tude and alt i tude rate) can be monitored by thecrew using a dual tape meter. Near hover, the twohorizontal components of velocity from the landingradar can be displayed on a cross-pointer meterdisplay.

**Actual ly, this is the t ime to reach the targetedcondit ions. The t ime remaining in the phase is 62set less than TG.

it spring back to the detent. Each side redesignationmoves the targeted landing site approximately 2.0deg; each forward or backward redesignation movesthe targeted landing site approximately 0.5 deg.When TG (Rl in Noun 64) goes to zero, the crewcan no longer redesignate by this technique. I f thec.rew d i d n o t h i n g a t t h i s p o i n t ( L O - G A T E ) , t h ecomputer would $utomatically select the auto-landingp r o g r a m ( P 6 5 ) . The crew normally switches outo f P 6 4 , h o w e v e r , and into the semiauto program(P66)-at about 500 feet. Program P66 is enabledby switching the spacecraft alt itude-mode switchfrom AUTO to ATTITUDE-HOLD. In P66, the crewcontrols the att i tude to maneuver the spacecraft toa desired landing s i te , but the engine thrott le iscontrol led by the computer to maintain a desiredalt i tude or a l t i tude rate. The crew contro ls thealt i tude rate by a rate-of-descent (ROD) switch,One act ivat ion of this switch changes the vert icalvelocity by 1 ft/sec in the direct ion se lected (up ordown). The crew can also control the engine throttlem.anually b y e n a b l i n g t h e f u l l y m a n u a l p r o g r a m(P67). To enable this program, the crewmemberswitches the thrott le-control switch from AUTO toM AI u AI,. To have manual att i tude contro l o f thespacecraft , he must select att i tude hold on theat.titude-mode switch, enabling manual attitude con-trol through the digital autopilot (DAP). Normally,P67 is not expected to be used, as P66 is a mucheasier mode to f ly the vehicle in, and i t appears towork well in f l ight.

From 500 feet on down, the crew selects the finallanding site very carefully to make sure that it isreasonably flat and free of boulders. Once they haveselected their f inal s i te , the crew nul ls the trans-lat ional velocit ies and reduces the vert ical velocityto 3 ft/sec, or less, by the t ime the contact l ightcomes on (al t i tude 5 feet ) . M’hen the contact lightdoes come on, the cre\v manually shut down the DPSengine. During this period, all systems, particularlyfuel remaining, are monitored very carefully by bothcrews-air and ground.

During the descent, secondary cues are alsomonitored; e.g., at 19,000 ft , observing the horizonin the forward window ( F i g . 17a); o b s e r v i n g t h elocat ion of the earth in the front windows (Fig.18);and, as mentioned earl ier , monitor ing the groundtrack while the windows are facing down (before40,000 ft).

Once safe ly on the ground, the crew cal ls thelanding confirmation program - (P68) . The objectof this program is to terminate the landing guidance,set the DAP funct ions, and init ia l ize the LGC forlunar surface operations. T h e s a m e p r o g r a mdisplays the computed posit ion for the landing site.Specifically, Noun 43 displays in Rl the Lat i tude,R2 the Longt i tude and R3 the AItitude o f the LMabove the lunar radius of the targeted landing si te .

Although I have l imited my discussion to theprimary system, there are comparable proceduresfor the backup, abort-guidance system (AGS). In

*Note: For P65, 66 and 67 the display changes to aNoun 60 (Fig. 15). Again, register one is the onlynew data. F o r N o u n 6 0 , Rl i s t h e h o r i z o n t a lcomponent of velocity.

9

passing, I should note that the backup system’sinertial package can be aligned to the same inertialorientation as the primary system; also that (by Verb47) the time and state vectors (CM and LM) storedin the primary system computers can be transferredto the backup-system (AGS) computer by using theLGC digital downlink.

IV MAN-MACHINE CONTROL INTERACTIONFOR SPACECRAFT ATTITUDE CONTROL

The description of the powered-descent phase ofthe lunar landing details man’s supervisory role inApollo. The levels of control interaction availableare illustrated in Fig. 20a and b, which summarizethe rotational attitude-command modes for thecommand-module primary and backup systems.

For the primary system, there are four spacecraftcontrol modes (Ref. 18): (1) automatic maneuverto a direction specified by either the computer orthe man; (2) attitude hold about a specified attitudewith selectable deadband and response rate; (3)manual maneuver with a three-axis hand controllerat a fixed, but selectable, rotation rate that is alsocoupled to the attitude-hold mode whenever thecontroller is returned to its neutral position; and(4) amode for small attitude maneuvers with timed,short, thrust impulses from the reaction-controlmotors, activated by a special three-axis controller.The latter mode is normally associated withcislunar-navigation observations.

For the backup system-it is really a number o:fsystems or parallel paths as Fig 20a and b illustrate--the spacecraft control modes, which are all manuallyactivated, are as follows: (1) an attitude hold withselectable deadband and response rate; (2) attitudemaneuvers with (a) the same three-axis hand control-.ler as the primary system and (b) a selectab1.emaneuver rate that is proportional to hand-,controller displacement; (3) attitude maneuvers withaccelerate command; and (4) attitude maneuvers,,with the normal controller, using minimum-thrustimpulses from the reaction-control motors.

Between the primary and backup systems, ninelevels of control interaction give four levels ofcontrol capability: (1) automaneuver; (2) manual.maneuvers with attitude hold; (3) manual maneuverswit’n timed minimum-thrust impulses, and (4) manual.maneuvers with accelerate command. The interfacesto these control capabilities range from direct handcontrollers to the computer DSKY. The necessityfor these nine levels of interaction was to providefunctional capability in the event of failures thatmight incapacitate large parts of systems. Again,there is great flexibility and redundancy, but heavyburden on the crew. As noted earlier, the increasesin reliability offered by the newer technology will.allow system structures that should not require thecrew tointeract with nine levels of control in orderto change the attitude of the spacecraft.

V SUMMARY

In summary, with the expected increases incomponent reliability and decreases in componentsize, future designs will provide airborne systems

with enough capability to relieve man of the necessityto play such an extensive role in either piloting orsupervising his vehicle.

For the time period of its design (1962- 64), themar-machine interface for the a.irborne computeroffers great flexibility for relatively small cost insize, weight, power, and computer memory. It shouldbe noted that the man can be integrated to this designin a variety of ways or levels. The lunar-landingsequence described earlier illustrates near-max-imum information flow and crew control of com-puter sequencing. The minimum level that couldbe implemented with this design would contain onlytwo programs, namely, POl, “Take me to the moon, ’and PO2 “Take me home,” The problem with thelatter mode is that it would be very inflexible anddifficult tomechanize on a development program aslarge as the first lunar-landing mission. The troublewith the present technique is that it does place asevere burden on the crew during training as wellas in flight. As lunar flight experience grows, abetter tradeoff between flexibility and complexity canbe made, even with the present equipment. Forexample, in cislunar space, the spacecraft’slongitudinal axis must be rotated slowly with respectto the sun in order to prevent large thermal gradients.This mode is called the passive-thermal-control(PTC) mode. During Apollo 8, this rate wasmaintained manually, by the crewman on watch, usingSCS minimum-impulse control. During Apollo 10and 11, revised procedures for initializing thelongitudinal rate, plus a new digital-autopilot (DAP)mode , removed the necessity for a crewman tocontinuously monitor the PTC mode. As a result,the crew can now all sleep at once.

Again, ne-wer technology offers control systemshaving variable-format, graphical input-output dis-plays. These newer systems offer easier (lessburdensome, hence less training), more-directfunctional communications. In the limit (Fig. 11,the computer, like Hal in the film 2001, can readman’s lips; or his mind as in The God Machine. Inmore practical directions, variable-format displayssolve the problem of panel space devoted to systemsused only for a few minutes out of the entire mission,e.g., systems for monitoring entry or boost.

Finally, these display/control systems, coupledwith a more administrative role for man-less directsupervising and sequencing as compared to Apollo-create the real challenge. The direction of thischallenge lies in determining the levels of interactionand the form that the interaction of man-machinecommunication should take. Apollo, therefore, isthe transition between (a) systems that require directconstant task interaction and supervision and (b)systems in which man-machine communication ismore functional, the supervisory role moreadministrative.

REFERENCES

1 Hopkins, A, L., Jr., A New Standard forInformation Processing Systems for MannedSpaceflight, MIT Instrumentation LaboratoryReport, R-646, September, 1969.

10

Hoag, D. G., Apollo Navigation, Guidance, andControl Systems- A progress r e p o r t , MYTInstrumentation Laboratory Report, E-2411,April, 1969.

Trageser, M. B., a n d H o a g , D . G . , A p o l l oSpacecraft Guidance System, MIT Instrum~tion Laboratory Report, R-495, June, 1965.

Hoag, D.G., Apollo Guidance and Navigation- AProblem in Man and Machine Integration, MITInstrumentation Laboratory Report, R-411, April1963, given at the AIAA conference in Dallas,Texas, April 22, 1963,

Dahlen, J. M., and Nevins, J. L., Navigation forthe Apollo Program, MIT InstrumentationLaboratory Report, R-477, May 1964, given atthe National Space Meeting of the Institute forNavigation, St. Petersburg, Florida, April 30,1964.

Sears, N. E., “Technical Development Status ofApollo Guidance and Navigation.’ Paoer 64- 178.Advances in the Astronautycal Sciences, pp. 114:450, Vol. 18, 1964.

T r a g e s e r , M . B . , a n d H o a g , D . G . , A p o l l oSpacecraft Guidance System, M I T I n s t r u -mentation Laboratory Report, R-495, June 1965,given at the IFAC -Symposium on “AutomaticControl in the Peaceful Uses of Space”.

Battin, Richard H., Explicit and Unified Methodsof Spacecraft Guidance Applied to a LunarMission, presented at the 15th InternationalAstronautical Congress, Warsaw, 1964.

Battin, R i c h a r d H., Astronautical Guidance.McGraw-Hill Book Co., Inc., New York, 1964.

11 Muller, E. S., and Kachmar, I?., The ApolloRendezvous System-Theory, Development andPerformance, MIT Instrumentation LaboratoryReport, R-649, 1969.

12 Klumpp, A. R., A Manually Retargeted AutomaticLanding System for LM bHT InstrumentationLaboratory Report, R-53; Rev. 1, August, 1967,

13 Battin, R. H., and Levine, G. M., Application ofKalman Filterinp Techniques to the ApolloProgram, MIT Instrumentation LaboratoryReport, E-2401, April 1969.

14 Nevins, J. L., Man - Machine Design for the ApolloNavigation, Guidance, and Control System,presented at the 2nd IFAC Symposium onAutomatic Control in Space, Vienna, Austria,September 4-8, 1967.

15 Nevins, J. L., Johnson, I. S., and Sheridan, T.B., Man/Machine Allocation in the ApolloNavigation, Guidance, and Control System, MITInstrumentation Laboratory Report, E-2305, July1968.

16 Duncan, R. C., “Man, Machine, and Informationin Flight Systems”, pp. 40-45, MIT TechnologyReview, February, 1969.

17 Alonso, R. L., AGC Keyboard and DisplayFormats, an internal MIT InstrumentationLaboratory Memorandum, Digital DevelopmentMemo No. 40, 26 September 1962.

18 Widnal, W. S., et.al., Guidance SystemsOperations Plan for Manned CM Earth Orbitaland Lunar Missions using Program Colossus 1- Section 3 Digital Autopilot, MIT Instru-mentation Laboratory Report, R-577, 1968.

10 Martin, F. H., and Battin, R. H., “Computer-Controlled Steering of the Apollo Spacecraft”,Journal of Spacecraft and Rockets, Vol. 5, No.4, 1968.

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3 0 External AV3 1 Lambert Aim Point Maneuver3 2 Co-Elliptic Sequence Initiation (CSI)3 3 Constant Delta Altitude (CDH)3 4 Transfer Phase Initiation (TPI)3 5 Transfer Phase Midcourse (TPM)3 8 Stable Orbit Rendezvous (SOR)3 9 Stable Orbit Midcourse (SOM)

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Fig. 8 Programs for Lunar Landing Mission

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E -2476

DISTRIBUTION LIST

D. HoagW. StamerisG. MayoN. SearsA. LaatsL. LarsonR . BattinE.C. HallJ. LawrenceG. StubbsR. CrispJ. HarperP. FellemanG. SilverJ . GilmoreG . OgletreeJ. BarkerT. Hemker (NAR)J. HandT. Lawton (MSC)M. JohnsonG. Cherrys . coppsM. HamiltonJ. Nevins (150)T. FitzgibbonR. O’Donnell (KSC)G. Levine

A P O L L O L i b r a r y ( 2 )MIT/IL Library (6)

R. ScholtenT. BrandB. KriegsmanE. Mullerc . P uP. PlenderJ . KernanJ. MorseA . EngelR. BairnsfatherJ . TurnbullR. SchlundtD. KeeneR. GossC. WorkE. JonesR. StengelD . F r a s e rE. OlssonR. MetzingerI. JohnsonJ . DunbarA . WoodinT. AndersonJ. O’ConnorR . McKernR. CooperD. BowlerE. Hall

J. FeldmanJ. CorriganG. B u k o wR. BoothT. ShuckA. SaltzmanJ . P a r rL. JohnsonL . Y o r g yG. KarthasC. GrayD. Y ankovich0. AndersonM. JohnstonR. LarsonB. McCoyM. SmithR. WhiteP. HeinemannM. AdamsR. LonesG . McWeeneyG. EdmondsJ. HarrisonG. SchmidtP. MimnoF . GlickA. HopkinsD. Hanley

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External:

AC ElectronicsKollsmanRaytheon

MSC:National Aeronautics and Space AdministrationManned Spacecraft CenterHouston, Texas 77058ATTN: Apollo Document Control Group (BM 86) (i8& 1~)

M. Holley (2)T. Gibson (1)

KSC :National Aeronautics and Space AdministrationJ. F. Kennedy Space CenterJ. F. Kennedy Space Center, Florida 32899ATTN: Technical Document Control Office

LRC:National Aeronautics and Space AdministrationLangley Research CenterHampton, VirginiaATTN: Mr. A.T. Mattson

GA:Grumman Aerospace CorporationData Operations and Services, Plant 25Bethpage, Long Island, New YorkA T T N . M r . E . S t e r n

NAR:North American Rockwell, Inc.Space Division122 14 Lakewood BoulevardDowney, California 90241ATTN: CSM Data Management

D/ 096-402 AE99

NAR RASPO:NASA Resident Apollo Spacecraft Program OfficeNorth American Rockwell, Inc.Space Division122 14 Lakewood BoulevardDowney, California 9024 1

GE RASPO:NASA Daytona Beach OperationsP. 0. Box 2500Daytona Beach Florida 32015A T T N : M r . H . L y m a n

(1)(3)(2)

(2)

(2 l& 1R)

(1R)

(2)

(3& 1R)

(8Pr 1 R)

(1)

(1)

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