apollo experience report guidance and control systems

8/8/2019 Apollo Experience Report Guidance and Control Systems

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N A S A T E C H N I C A L N O T E NASA TN D-8249

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APOLLO EXPERIENCE REPORT -GUIDANCE AND CONTROL SYSTEMSRaymond E. Wilson, Jr.Lylzdon B. Johizson Space CenterHouston, Texas 77058N A T I O N A L A E R O N A U T I C S A N D S PA CE A D M I N I S T R A T I O N W A S H I N G T O N , D. C . JUNE 1 9 7 6


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1. Report No.NASA TN D-8249APOLLO EXPERIENCE REPORTGUIDANCE AND CONTROL SYSTEMS

2. Government Accession No. 3. Recipient's Catalog No.

I June 19764. Title and Subtitle6. Performing Organization CodeI JSC-086305. Report Date

7. Author(s)Raymond E. Wilson, Jr.9. Performing Organization Name and Address

Lyndon B. Johnson Space CenterHouston, Texas 77058

8. Performing Organization Report No.JSC S-41810 . Work Unit No.914-50-00-00-7211 . Contract or Grant No.

2. Sponsoring Agency Name and Address13 . Type of Report and Period Covered

Technical NoteNational Aeronautics and Space AdministrationWashington, D. C. 20546

7. Key Words (Suggested by Author(s))'Navigation ' Flight Control' Stabilization Contr ol 'Rendezvous' Attitude Control ' Reentry Guidance' Reaction Control Gyroscopes* Descent Trajectorie s

14. Sponsoring Agency Code

18. Distribution Statement

Star Subject Category:12 (Astronautics, General)

I5. Supplementary Notes

19. Security Classif. (o f this report) 20.Security Classif. (of this page) 21 . NO . of PagesUnclassified Unclassified 20

6. AbstractThe Apollo guidance and control sys tem s fo r both th e command module and the lunar module aredescribed in a summary report. General functional requi rement s are discussed, and generalfunctional descriptions of the various subsystems and their interfaces are provided. Thedifferenc es between the or iginal in-flight maintenance concept and the final lunar-orbita l-rendezvous concept are discussed, and the background in philosophy, the syst em development,and the re aso ns fo r the change in concept are chronologically presented.the command module guidance and cont rol sys te m under each concept ar e included. Significantconclusions and recommendations contained in more detailed reports on specific areas of theguidance and control systems are included.

Block di ag ra ms showing

22 . Price'$3.50

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CONTENTS

Section

SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .DESCRIPTION O F SYSTEMS . . . . . . . . . . . . . . . . . . . . . . . . . .

Command and Se rvi ce Module . . . . . . . . . . . . . . . . . . . . . . . . . .Lunar Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

CONCLUDING REMARKS AND RECOMMENDATIONS . . . . . . . . . . . . .REFERENCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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FIGURES

Figure1

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910

PageDiagram of Block I command module guidance and

control syste m . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4Diagram of Block 11 command module guidance and. . . . . . . . . . . . . . . . . . . . . . . . . . . . 5ontrol systemCommand module primary guidance, navigation, and 6ontrol system . . . . . . . . . . . . . . . . . . . . . . . . . . .Command module sta biliza tion and control sy st em . . . . . . . . . . 8Entry monitor s yste m . . . . . . . . . . . . . . . . . . . . . . . . . 9The ORDEAL ass embl y, showing panel and switch positi ons . . . . 9Lunar module guidance and control sy st em 10onfiguration di agra ms . . . . . . . . . . . . . . . . . . . . . . .Lunar module pr ima ry guidance, navigation, andcontrol system . . . . . . . . . . . . . . . . . . . . . . . . . . . 11Lunar module stabilization and control syst em . . . . . . . . . . . . 12Lunar module abort guidance system . . . . . . . . . . . . . . . . . 12

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APOLLO EXPER IENCE REPORTGUIDAN CE AND CONTROL SYSTEM S

B y R a y m o n d E. W i l s o n , Jr.Lyndon B. J o h n s o n S pa ce C e n t e rS U M M A R Y

This r epo rt contains gene ral descri pti ons of the Apollo guidance and cont rolsy st em s fo r the command module and the lunar module and is the general summarydocument fo r a se ri es of more detailed repor ts on specific areas of the guidance andcontrol syste ms. The Apollo Block I guidance and control system is briefly describedto show the dif fer ence s between the orig inal in-flight maintenance concept and the fina llunar- orbita l-rende zvous concept. Even with the in-flight maintenance concept, thereliability of the Block I design was l es s than that of the Block I1 design, which incor-porated built-in redundancy and al ter na te modes of operation . The Block I1 conceptincludes sev era l unique sys tem s that provide the nec essa ry prim ary operational func-tion s and the nec ess ary redundancy fo r backup and abor t requir emen ts. Some of themore significant recommend ations and conclusions contained in the other Apollo guid-ance and control system re port s are included i n this summary report.

INTRODUCTIONThe Apollo guidance and control (G&C) ystems for the command and se rvic emodule (CSM) and the lunar module (LM) were designed and qualified to support the14-day lunar mission. (Information regard ing the contr ol of unmanned Apollo mis sions

is given in refs. 1 and 2.) The design of t he G&C sy st em s included the following sub-syst ems: the stabilization and control system (SCS); the primar y guidance, navigation,and control system (PGNCS); the e ntry monitor syst em (EMS);he orbital-rate drive,Earth ana lunar (ORDEAL) assembiy ; he service propulsion sy stem (SPS) imbalactuators; the mission control programer (used on the Block I command module (CM)unmanned flights); and the abort guidance system (AGS).

The Block I CM des ign philosophy wa s based on an in-flight maintenance concept.Fro m e ar ly in the Apollo Pr ogra m until the major Block I1 CM change in June 1964,the SCS w a s the primary method of flight control. The initial Apollo proposal indicatedthat the c ontrol system would al so encompass a luna r landing capability. This idea,NASA. A s the pr im ar y method of flight control, the Block I system had to meet ahowever, wa s eliminated when the lunar-orbital-rendezvous concept was adopted bystandar d of high relia bility th at could be achieved only through the use of in-flight


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maintenance. Despite the best ef fo rt s of reliability and design engineers, the meanti me between fai lu re s fo r the equipment wa s of the sa me o rd er of magnitude as themission t ime. Hence, the mos t fe as ib le solution to the problem of maintaining highsys tem reliability w a s to provide standby redundancy in the fo rm of onboard s pa re s.

A t the tim e the Block I1 concept was being defined in June 1964, the in-flightmaintenance requi rem ent was abandoned. Becau se the Block I Earth-orbital reliabilityrequirement was considerably less strin gent than that fo r the Block I1 lunar mission,additional redundant c ir cu it s we re not added to the Block I vehicle in l i e u of onboardsp ar es . Instead, the Block I1 system was re designed to include redundant con trol paths.

The concept of in-flight maintenance wa s disca rde d becaus e it proved to beimpractical. Although it was technically f easibl e fo r the astronau t to detect and repl acea defective module, it w a s not an eas y task. Fo r example, in the control system lab-ora tor y, experienced and highly trained tech nicia ns often requi red many hou rs to locateand change a defective element. Furthe rmore, the changes necess ary for meeting anew requir ement to withstand much gre at er humidity made the install ation and remo valof the modules even mo re difficult.

In the sum me r of 1964, t he Apollo P ro gr am wa s divided into the Block I andBlock I1 vehicle p rog ram s; the difference was that the Block I1 veh icl es would havelunar -miss ion capability. The underlying concept of th is change was that the PGNCSshould be considered the p ri ma ry mode of con tro l and the SCS the backup mode; that is,the SCS was to be used f or contr ol when the PGNCS wa s not used.

The Block I1 design of built-in redundancy and al te rnate modes of ope rat ion alcapability has resulted in a G&C sys tem that is more reliable and that weighs less thanthe Block I system. In addition, the stand ard conn ect ors used on Block I1 have beenless troub lesom e than the conn ector s used on Block I. The pro ble m of single-pointfa i lures wi l l lim it the reliabili ty of any sys tem . The complexity of spa ce sy st em sprodu ces a maze of fai lure modes and subtle cir cu it s that are almost impossible toanaly ze by "brute for ce" methods. Analy sis too ls and metho ds are needed to assist inthe tota l design pro ces s and t o ensu re the operation of the systerh.

The init ial concept and configuration of the LM G&C syst em evolved dur ing 1963and 1964. The initi al concept wa s that the Government-furnished pr im ar y guidancesy st em would provide the n ec es sa ry guidance and navigation (G&N) functions and thecontractor-furnished SCS would provide the vehicle stabi lizat ion and con tro l functions.In addition, the SCS was t o provide a backup guidance capability sufficient to p er mi tinsert ion into a safe lunar orbi t if pri mar y guidance were l ost.

By mid-1964, the stabi lizat ion and co ntr ol functions had become fa ir ly welldefined, and desig n-con trol spec ifica tions had been completed and su bco ntr act s award edfo r most assemblies. During the fall and winter of 1964, NASA (with contractor partic-ipation) reviewed the LM G&C req uir eme nts and the ha rdware capabi lities of the pr i-ma ry guidance sy st em and the SCS. Th is review resu lte d in implementation of theintegrated G&C concept. In addition, the backup or a bor t guidance r equ ire men ts provedto be more complex than orig inall y envisioned.

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The LM G&C sys tem incl udes two partia l paths o r sy st em s for vehicle guidanceand control. The PGNCS prov ides the nec ess ary G&C capability for mi ssi on comple-tion. The AGS provides the nec ess ary G&C capability fo r missi on abo rt in the event ofPGNCS fai lur e, but does not provide a missio n completion capability. Such a capabilitywould have required a backup system having the same capability as the PGNCS and wasprohibitive in cos t, weight, and power. The SCS fo rm s an in te gr al pa rt of both theprimary and abort systems. A s par t of the primary syst em, the SCS includes thedri ver s fo r reaction control system (RCS) jet operation; the electronic interface fo rdescent-engine th ru st and gimbal cont rol; and the hand c on tr ol le rs fo r manual-attitude,desce nt-th rust , and tran slat ional input commands. In the abor t system, the SCS pro-vides jet-select logic, signal summing, and gain cont rol, and the hand con tr ol le rsused fo r manual input commands are the sa me as those used in the pri mary system.Atti tude reference o r s teering er ro rs are provided to t he SCS by the AGS.

The development of the PGNCS digital autopilot is described in reference 3 . Thedevelopment of the G&C hybrid simulation facility by the LM co nt ra ct or is described inreference 4.

D E S C R I P T I O N OF SYSTEMSThe CSM and LM G&C systems are described in the following paragrap hs.

Comm and and Serv ice ModuleA brief functional des cri pti on of the CSM G&C sy st em s follows. The relationsh ip

of the SCS and the PGNCS for the Block I and Block I1 vehicles is shown in figures 1and 2, respectively.The PGNCS is Government-furnished equipment (GFE) nd is common to both theCSM and the LM except for diff erenc es in the optics, minor di ffere nces in the ine rti al

measu remen t unit (IMU) (acce lerom eter scaling and location of acc ele rom ete r elec-tr on ic s components), and differen ces in the computer software programing. The sedifferences are described in gr eat er detail in reference 5. The PGNCS is divided intothree major subsystems- nert ial, optical, and computer- nd designed so thateach subsystem can be oper ated independently during a n eme rgen cy o r backup mode.Therefore, the fa il ure of one sub sys tem will not disable the en ti re PGNCS. Thethree subsystem-s, o r combinations of them, can perform the fc!!owir,g functions.

1. Periodically establis h an inertia l reference used f or measu rement s andcomputations2. Calculate the position and velocity of the sp ac ec ra ft by opt ica l navigation andine rti al guidance3 . Generate steer ing signals and calculate targeting data and thrus t commands

nec ess ary to maintain the required spacecraft trajectory4. Provide the astronaut with a dat a display that indicates the sta tus of the G&N

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Display andS e d a n tc a n n i n gtelescope keyboardIin k

Figure 1. - Diagram of Block I command module guidance and cont rol sys te m.

Display andkeyboard

The PGNCS equipment (fig. 3 and refs. 5 and 6 ) con sis ts of a navigation base, anIMU, an optical assembl y, a power and s er vo assembl y (PSA), an Apollo CM computer(CMC), display and control panels, and an elec tro nic coupling data unit. The navigationbase is mounted t o the spa cec raf t sidewall and is used as a holding fix ture f or the IM Uand the optical assembly. The IMU and the op tical assem bly are attached to and p re -cisely alined with the navigation base. The display and control panels co mpr ise thefront of the G&N struc ture and are located so that an as tron aut can view and manuallyopera te them. The PSA is located on a shelf below the navigation base. The electroniccoupling data unit and the CMC are located on a shelf below the PSA. Two display andkeyboard (DSKY) nits ar e located in the vehicle, one at the ma in panel and the othe r inthe lower equipment bay. The DSKY provides, access to the CMC and furnishe s statusinformation t o the ast rona uts.

The inerti al subsystem (ISS) cons is ts of the IMU, the el ectro nic coupling dataunit, and portions of the PSA and the display and cont ro l pan els . The IS S measureschang es in space cra ft attitude, assists in generating steering commands, and mea sur esspa cec raf t velocity changes. The IM U provides an iner tial refere nce consisting of astable member gimbaled fo r th re e deg re es of fr eed om and stabilized by signals f romthr ee integrating gyros. Acceleration of the spacecr aft is sensed by t hr ee pendulousacc ele rom ete rs mounted on the stable membe r and having orthogonal input axes. Theresultant signals fro m the ac celero meters are supplied to the CMC, which then calcu-lates the tota l velocity. The modes of opera tio n of the IS S can be initiated ei the r auto-matically by the CMC o r manually by the ast ronau t.4


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-Display and Display and. teyboard keyboardextantpdata I Scanningtelescopein k I

IJ

State vector I II -Voice I Inertial Commandcommuni- - measurement - odulecations un i t computer - Je t on-off signals-- -- - - - -rimary navigation Primary guidance and controlI contro l system_ - - _ _ _ _ _ L---'------Astronaut Stabilizationdisplays system I SPSgimbal -ervomechanismsCS attitudereference and

s PSgimbal -I - servomechanismsSCS contro ls I

IFigure 2. - Diagram of Block I1 command module guidance and c ont rol sy ste m.

The optical subsystem (OSS) ons ist s of a sextant and telescope, the electro niccoupling data unit, and port ions of the PSA and the display and co ntr ol panels. The OSSpro vid es the CMC with da ta obtained by measur ing ang les between l ines of sigh t tocelestial objects and provides measur ements for establishing the in ertia l reference.The sextant is a high-magnification (284 , dual-line-of -sight device used for preciseangular measurements. The telescope has a wide fie ld of view and one line of sight andis used fo r co ar se acquisition o r orbital tracking of landm arks. A manual controlstick is manipulated to position the optical lines of sight. A manually initiated timingma rk c au se s the CMC to r eco rd both the angle and the tim e at the insta nt the sextantis prop erly pointed fo r a measurement.

The computer subsystem consists of the CMC and po rt ions of the displa y andcon tro l pane ls. The CMC, which is used to perform space-flight da ta handling andcomputations, is a general-purpo se digit al computer consi sting of a cor e memory,paral lel operations, and a built-in self-check capability. Pr og ra ms sto red in the CMCare sele cted to control and solve flight equations. The selection of the pr og ra ms canbe controlled eith er manually o r by automatic sequencing. The computer sub syste mcalcu lates the steering sig nals and the engine discrete commands nec ess ary to keepthe spacecraft on a requi red traj ector y, positions the stable mem ber in the IMU to acoordinate sy ste m defined by preci se optical measurements, pe rf or ms limited mal-function isolation, and suppli es pert inen t spacecraft-condition information to the displayand control panels.

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NOTE:

Figure 3. - Command module pr im ar y guidance, navigation, and cont rol sy st em .The SCS (fig. 4 and ref. 7 ) provide s (to the PGNCS) backup sta bil iza tion and con-tr ol of the spa cecr aft fo r rotational , transl ationa l, and SPS thrust ing, using the CSMRCS and SPS-engine gimbal se rvo mec hanisms . The SCS al so provides the display s andcontrol s required for crew interface. The SCS is divided into thre e basic subsyst ems:attitude reference, attitude control, and thrust -vector control. The subs yste mscontain the elem ents that provide sele ctable functions fo r displa y, automat ic and

manual attitude control, and thrust -ve ctor c ontrol . The components and functionsof the SCS hardware are as follows.

1. The reaction -jet and engine on-off cont rol contain the solenoid dr i ve rs andthe logic circuits ne cess ary t o control the reaction-jet automatic solenoid coi ls and theSPS-engine solenoid control valves and relays.2. The electronic control assembly contains the circuit elements required forsumming, shaping, and switching the rate and attitu de-e rror signal s and the manual

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input s ignal s nec ess ary to maintain backup stabilization and cont rol in all axes fo rthrust-vector and attitude control.3 . The elect ronic display assembly provides the logic cir cu it s fo r establishingthe signal sourc es to be displayed and the displays to be used.4. The gyro disp lays coupler provides the inte rface between the body rate sensorsand displays to give an a ccura te readout of spacecraft attitude relative to a given refer-

ence coordinate system.5. The thrust-vector position servoamplifier provides the ele ctric al interfacesbetween the command el ec tr on ics and the gimbal act uat or fo r positioning the SPS engine.6 . Each of the two gyro as sem bli es contains thre e sensi ng eleme nts, body-mounted attitude gyros, and the ele ctro nics components ne ces sar y to provide output sig-na ls propor tional to angular rate or to angula r displacement fo r eac h of the thr ee bodyaxes.7. Two flight di re ct or attitude indicators display spac ecraf t attitude, attitude-er ro r, and angular rate information to the crewmen.8. The gimbal position and fuel-p ressu re indicator provi des a redundant displayof t he SPS-engine pitch and yaw gimbal angles and a mean s of introducing manual t r imof the engine gimbals. The indicator ha s the alterna te capability f or providing a displayof Saturn II and Saturn IV-B fue l and oxidizer pr ess ure s.9. The attitude- set control panel provides a means of manually estab lishi ng anattitude reference coordinate system and a visual readout of the coo rdi nat es commanded.

10. The translational controller provides a mea ns of ex erci sing manual controlove r re cti lin ear motion of the space craft in both direc tion s along the thr ee pri ncipalbody axes. The translational controll er also provides the capability f or manual abortinitiation during launch by counterclockwise rotation and for tra ns fe r of s pac ecr aftcon tro l fr om the PGNCS to the SCS by clockwise rotation.

11. Two rotational co ntr oll ers provide a means of exerc ising manual control ofspace craft rotation in eith er direction about the thr ee main axes and pr ovides formanual thrust-v ecto r contr ol in the pitch and yaw axes.

The E M S (fig. 5 and ref. 8) prov ides information tinat enabies the crewmen tomonitor the PGNCS-controlled ent ry performance and velocity changes, to provideth ru st term inati on sign als under SCS-controlled velocity chang es, to manually contr olentry if the PGNCS fails, and to display very-high-frequency rang ing informationobtained between the undocked CM and LM . The sys tem inc ludes both har dware andsoftware. The software aspec ts refer to the generation of flight-patte rn lim it lines.The EMS con sis ts of two basic a sse mbl ies : the entry monitor cont rol assemb ly (EMCA)and the entry monitor s cr ol l assembly (EMSA). The EMCA contains the elec troni cscomponents and is made up of integ rated cir cuit s, a range indicator, an accelerometer,the velocity-change/range-to-go counter logic, a pulse scaler, power supplies, andswitches. The EMSA or load-factor/velocity (g-V) plot ter assembly cons is ts of asc ro ll of M ylar tape o r fil m impri nted with g-onset, g-offset, and ran ge potential lin es.

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The EMS display co mpr ise s fou r functional components ess ent ial to tr aje cto ry monitor-ing and flight control: the ro ll attitude indicator, the ent ry threshold in dicator, theco rr id or verification indicator , and the flight monitor o r scroll .

Electronic controlassembly

R e a c t i o n - j e t a n d e n g i n eon-off control

Electronic displayassembly Gyro displayc o u p l e r Gyro assembl ies 1 an d 2

Figure 4. - Command module sta bili zation and cont rol syste m.The rol l attitude indicator disp lays the angu lar position of the lift vect or about therel ati ve wind vect or of the vehicle. The en try thre shold indi cato r is a lamp that illumi-nat es when the vehicle e ncounters a threshold aerodynamic acceleration level, normallyan acceleration load fact or of 0.05g. The co rr id or verificatio n indicator con sis ts of twolamps, one of which illuminates at a prescribed time (normally 10 seconds) after theentr y threshold is reached. The particu lar lam p illuminated dep ends on the mea sure daccelerati on level. The flight monitor or scr ol l is the major component and provid esa rectilinear presentati on of the entr y-ac cele rati on load fact or (g) as a function of iner-tial velocity (V). The display is created with a Mylar tape that ha s a monitoring pattern

printed on the for mat and an emulsion bonded on the back. The ver tic al axis is drivenin proportion to the accel erat ion load fac tor . A scribe remov es the emulsion to createthe entry trajectory g-V t ra ce as the tape translates horizonta lly; the flight cre w thencompares the trace to the permanently displayed monit oring pattern.

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Functionselectorsw itch-,

S cro l lentrypatterns

LLower corridorannunc iat or num er ic d isplay

Figure 5. - Entry monitor system. Figure 6. - The ORDEAL assembly ,showing panel and switch positions.Th e ORDEAL (fig. 6 and ref. 9) is a device that can be switched dire ctly into theFDA1 total-attitude cir cui ts to perform a coordinate trans forma tion of space craft pitch

from an inertial to a local-vertical refer ence frame . The ORDEAL assem bly is sup-plied as GFE and is comm on to both the CM and the LM. When the sp ac ec ra ft panelswitches are in the orbita l-rat e position, si ne s and cosi nes of the pitch angle are appliedas inputs to a pair of reso lv er s in the ORDEAL ele ctr ome cha nic al module. When thespacecraft panel switches are in the no rm al position, the ORDEAL is bypassed and theindic ators display pitch attitudes to the in ertia l fr am e of refer ence.

Lunar ModuleA brief functional description of the LM G&C sy st em follows. The functionalrela tions hip between the sy st em s that make up the G&C sys tem is shown in figur e 7.The PGNCS (fig. 8) se rves as the autopilot in con trol ling the LM throughout themission. Norma l guidance require ments include tra nsf err ing the LM fro m a lunar orbit

to the des cen t profile, achieving a succe ssfu l landing at a preselected o r crew-selectedsite, and performing a powered ascent and rendezvous maneuver that results in ter-mina l rendezv ous with the CSM. If the mission is to be aborted, the PGNCS pe rf or msguidance man euv ers that pla ce the LM in a parking orbit o r in a trajectory that inter-cep ts the CSM tra jec tor y, The LM missior, p r o g r a ~ e rs described in rzference 10.

The PGNCS includes th re e maj or subsyst ems: inert ial, optical, and computer.The ISS establis hes the ine rti al reference fra me that is used as the centra l coordinatesyste m from which al l measurements and computations are made. The ISS mea su re sattitude and inc rem ent al velocity ch anges and assists in converting dat a for computeruse, onboard display, o r telemetry. Operation is sta rte d automati cally by the guidancecomputer or by an a stron aut using the computer keyboard. Once the syste m is ener-gized and alined to the inert ial refe ren ce, any LM rotation is sense d by the stabl e mem-ber . The alinement optical telescope, a unity-power, peri scope -type dev ice with a

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60" conical field of view, is operated manually by the astronauts, and data are read outand manually inse rted into the compu ter through the DSKY. Th is fe atur e is the majordif ferenc e between the CSM PGNCS and the LM PGNCS.

radarLandingradar

Manual inputs1 I

commandsand altitude Primary guidance,navigation, andcontrol system 1

Communications Position, velocity, time, attitude1 1

II

ISelected on-off orI-*+ handover commands

guidancesystem

AbortdataStabilizationand control

IAutomatic manualattitude and translationalautomatic propulsion -

On-off, t rim, and thrott le commandsLLA-Figure 7. - Lunar module guidance and control system configuration diagrams.

The computer syst em, as the control and data -proc essin g c ent er of the LM,enab les all the G&N functions neces sar y f or automatic c ontr ol of the path and attitudeof the vehicle. The LM comput er is identical to the CSM computer except fo r theabs enc e of stor ed pro gra ms.

The SCS pro ces ses RCS and main propulsion sys tem c ontr ol signals for vehiclestabilizat ion and cont ro l (fig. 9 and ref. 11). To stab ilize the LM durin g all pha ses ofthe mission, the SCS provides signa ls that fire proper combinations of 16 RCS thrust-ers. These signals control attitude and trans latio n about o r along all axes; data inputsorigi nate from ei ther the PGNCS o r the AGS. The SCS al so pro ce ss es on and off com-mands for the ascent and descent engines and rou tes automatic and manual throttlecommands to the descent engine. Tr im c ontro l of the gimbaled desc ent engine is alsoprovided by means of gimbal drive actuators, to ensu re that the thrust vector pa sse sthrough the LM cent er of grav ity. The SCS cons is ts of two atti tude co nt ro ll er as se m-bli es (ACA), two thr ust and tran slati on cont rol ler as se mb li es (TTCA), an attitude andtranslation contr ol asse mbly (ATCA), a rate gyr o ass emb ly (RGA), gimbal dr iv eactuator (GDA), and a descent-engine c ont rol ass emb ly (DECA).

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Power an d servo

Figure 8. - Lunar module pr im ar y guidance, navigation, and con trol sys tem .The ACA supply attitude rate commands (proportional to t he dis placement of the

sti ck) to the computer and to the ATCA, supply an out-of-detent discrete command eac htime the handle is out of its neut ra l position, and supply a followup di sc re te commandto the AGS.Functionally, the TTCA are three-axis integrated translation and thrust control-lers that enable astronaut s to command vehicle translations by fir ing RCS th ru st er s

and t o thro ttle the desc ent engine between 10- and 92.5-percent th ru st leve ls.The ATCA contr ols LM attitude and translation. In the FGNCS path, attitude andtranslational commands are generated and applied directly to jet dr iv er s within theasse mbly . In the AGS path, the ATCA receiv es tran slat iona l commands fr om the

TTCA, rate-dam ping sign als fro m the RGA, and attitude/rate comma nds and pulsecomma nds fr om the ACA. The asse mbly combines attitude and tran slat iona l commandsin its logic network to select the proper thrusters to be fire d fo r th e combination oftr an sl at io n and rotation. The DECA acc ept s engine-on and engine-off com man ds fr om

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the SCS contr ol assem blies, throttle com-man ds fr om the PGNCS and the TTCA, andtri m commands from the PGNCS o r thecommands to the desce nt engine and rou testrim commands to the GDA.

A t t l t u d e a n d

assembly ATCA. The ass emb ly appli es throt tlet ranslat ionc o n t r o l

Th e AGS (fig. 10 and ref. 12), whichc o n t r o l a s s e m b ly is used as a backup for the PGNCS duringan LM missi on abort, det erm ine s the LMtrajectory o r t ra ject or ies required forrende zvous with the CSM and ca n guide theLM fr om any point in the mission , fr om

Rate gyro assembly CSM/LM sep ara tio n to CSM/LM ren dez -vous and docking, including as ce nt fr om

assembly the lunar surfac e. The AGS can providedata fo r disp lays and for explicit guidancecomputations and ca n al so enable engineignition and shutdown.

assembly

The AGS con si st s of a nongimbaledinerti al refere nce package (the abort sen-sor assembly) that is rigidly strapped tothe LM rather than mounted on a gimbaled,stabilized platform. The abort sen sorassemb ly contains thr ee floated, pulse-

rebal anced , single-degree -of -freedom, rate-integrating gy ro s and th re e pendulousref ere nce acce ler ome ter s. The data entry and display assembly consists of a control

Figure 9. - Lunar module stabilizationand control syste m.

Figure IO. - Lunar module abort guidance system.1 2


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panel (to which electrolumine scent disp lays and da ta e ntry pushbuttons are mounted) anda logic enclo sure that hous es logic and input/output circu its. The abo rt elect ron icsassembly is a high-speed, gene ral-p urpose computer with spec ial-p urpos e input/outputele ctr oni cs components. The computer ha s a memory capacity of 4096 words, of whichhalf are permanent and half are temporary.The ORDEAL assembly in the LM is identical to the unit used in the CSM.

C O N C L U D I N G R E M A R K S A N D R E C O M M E N D A T I O N SDur ing the cou rs e of the development, qualification, and flight programs, theApollo guidance and contro l sy st em s performed in an outstanding manne r. The re wer eno guidance and control fai lu re s or malfunctions that precluded missi on completion o rthat placed the flight cr ew o r the missi on in jeopardy.In general, the approaches that were used to est abl ish and impleme nt guidance andcontr ol syst em interf aces and checkout procedures during the integra tion of the sy st emsin the spacecra ft appear to have been sound. Consequently, few inter face pro ble msappeared during the integration of the sys te ms into the spacecraf t. Some of the moresignificant it ems that dese rve ca ref ul consideration on futur e pro gr ams are as follows.

1 . A stro ng effort should be made t o establish baseline req uireme nts before thestart of har dware design and software development pr oces ses. Fo r example, changesaffecting hand con tro lle rs, humidity, and in-flight maintenance caused maj or redesignefforts.2 . A fai lur e-a nalysi s technique should be developed to assist in the identification

of single-point failures. The Apollo method, i n which many engineers must searchdiagr ams for problems, is not altogeth er successful for complex sys tems .3 . Criteria and methods to obtain extended-duration hardware reliability shouldbe establish ed. The long checkout ti me s and the extended dur ati ons of som e missio ns

put a premium on checkout, storage, and long-life operation of equipment.4. Seri ous considerat ion should be given to the use of solid-state devic es fo r

switching inst ead of r el ay s and switch es. If relays are used, superscreening testsshould be esta blished to ens ure high reliability.5. System integration can be best achieved if a singl e vendor sup pli es the hard.-

ware f or system-level requirements such as the command and se rv ice module stabili-zation and control system. Th is approach cont rasts with the lu nar module stabilizationand control syst em procurement, which wa s at the assembly level with the contractorretaining responsibility f o r integration. Although system -leve l procureme nt is not apanacea fo r all problems, inter face problems are mor e eas ily avoided than withassembly-level procurement.

6. Thermal-vacuum testing revealed as many failures as vibration testingrevealed. Both envi ronments should be used in assembly acceptance testing.

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7. Numerous experien ces in the prog ram have demonstrated the nec essi ty fo rhaving immediate ac ce ss to an al terna te sou rce of qualified pa rt s to supp ort scheduledpro gra m milestones.8. Because of schedule cons tra int s, t he abort guidance sys tem progr am (as anexample) did not include the use of an engineering model fo r design evaluation befo re

the system w a s committed to production. Many manufacturing and circuit-d esign prob-le ms could have been cor rec ted e ar ly if an eng inee ring evaluation model had beenavailable.9. The initial integration of the guidance and contr ol sys te ms with the spa cec raf tcaused many pro ble ms durin g vehicle checkout. The participation of the subco ntr actors

(for example, tho se who built the individual pa rt s of the lunar module stabilization andcon tro l system) would have been valuable dur ing the reso lution of those prob lems. Areview by the subcontractor of vehic le checkout pr ocedur es and onsite sup port duringspa cec raf t integration would prove beneficial in future progr ams .

10. Definition of vehicle test methods, parti cular ly combined-systems andinteg rate d-sys tems testing, should be established ear ly to avoid inte rface and inter fer -ence problems.

11. Exper ience gained dur ing the development of the Apollo digital autopilots maybe used to avoid fut ure design problems. Logical deci sion techniques should be appliedwith care in design development because conditions may exist in which these techniquescan unexpectedly lock out en ti re system functions . The us e of logic in avoidingdegraded performance ha s to be traded off with unintended res tri cti ons .12. Fur the r re se ar ch effort should be expended to develop additional analyticaltechniques fo r digi tal cont rol syst em design. Adaptive design techniques making use ofthe inherent flexibility available in digital s ys te ms should al so be established.13. Design requirem ents should include the requi remen t to pr es er ve the capa-bility fo r monitoring sys tem effectiveness. Fo r example, efficient use of the del ta -

velocity capability of the Apollo ser vice propulsion sy st em placed only mild co ns tr ai nt son maintaining sm al l vehicle-attitude e r r o r s and rates during the start transient;however, a design goal w a s to produce a syste m that minimized these start transientsfo r nominal operati on so that the tr ans ien ts would be usefu l ind ica tor s of potentiallyserious off-nominal conditions.14 . Caution must also be exerc ised in placing too much relian ce on simulationresults for design verification without a full appreciation of the approximations thathave been made in developing the p ro ce ss m odels and in implementing th ese models in

the simulations. The implementation bec omes particula rly important when simulatinghigh frequency dyna mics in a digital computer.15. Engineering simulation should be recognized as a potentially large, expensiveoperation and, as such, should be given appro priat e attention during the ini tial contractdefinition and negotiations to ensure the est abl ish ment of well-defined base line p lansand costs.

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16. The ro le that simulation is expected to have in the program should be definedin enough specific det ail to enable establishment of an appr opr iat e simulation plan thatis adequate, but not unnecessarily elaborate. If possi ble, th is definition should beincluded in the request f o r proposal so that the con tra cto r can provide appr opria te plansand costs.

17. Management of the simulation activity should be delegated in so me appro-pri ate way so that the activity will receive adequate full-time attention.18. Consideration should be given to having la rge simul ators that r equire systemhardware constructed at Government facilities. Contractor personnel would be used as

requir ed duri ng construction and during initial phases of operation; civil s erv icepersonnel would be used later in the program.19. Detailed planning for large simulators should begin early in the program, butactual implementation should be delayed as long as possible t o avoid tr acking and incor -porating inter im ch anges to the s ystem being simulated.20. Detailed planning should be designed to ensu re the inclusion of req uir eme ntsfo r special-purpose equipment needed fo r interface or sys tem simulation, req uir e-ments for ext ern al scene generat ors, and provisions fo r dat a input and output. Carefu lplanning is important because these requir ement s can become expensive.2 1 . At the beginning of the proposed program, the degree of desired formalityassociated with the simulator operation should be determined so that proper plans canbe made. Configuration con tro l, documentation, extra sets of hard-copy data, formal

test-readiness reviews, and anomaly reporting can create a greatl y incr eased workloadfor support personnel.

Lyndon B. Johnson Space CenterNational Aerona utics and Space AdministrationHouston, Texas, November 8, 1974914-50-00-00-72

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REFERENCES

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11.

12.

Holloway, Gene F. : Apollo Experience Report- Guidance and Control Systems:Automated Control System for Unmanned Mission AS- 201. NASA TN D- 7991,1975.

Holloway, Gene F. : Apollo Experience Report - Guidance and Control Systems:Mission Control Programer for Unmanned Miss ions AS- 202, Apollo 4, andApollo 6. NASA TN D-7992, 1975.

Peters, William H. ; and Cox, Kenneth J. : Apollo Experience Report - Guidanceand Control Systems: Digital Autopilot Design Development. NASA TN D- 7289,1973.

Gilbert, David W. : Apollo Experience Report - Guidance and Control Systems:Engineering Simulation Pr og ra m. NASA T N D- 7287, 1973.Holley, M. D. ; Swingle, W. L. ; et al. : Apollo Experience Report - Guidanceand Control Systems: Pr im ar y Guidance, Navigation, and Control SystemDevelopment. NASA TN D- 8227, 1976.McMahon, William A. : Apollo Experience Report - Guidance and Contro l

Systems: CSM Service Propul sion System Gimbal Actuat ors . NASA TND- 7969, 1975.Littleton, Orval P. : Apollo Experience Report - Guidance and Cont rol System s:Command and Ser vice Module Stabiliza tion and Con tro l System. NASA TN

D-7785, 1974.Reina, Ben, Jr . ; and Patterson, Herber t G. : Apollo Experience Report -Guidance and Control Systems: Command and Se rv ice Module Entr y MonitorSubsystem. NASA TN D-7859, 1975.Parker, Roy B. ; and Sollock, Paul E. : Apollo Experience Report - Guidanceand Control Systems: Orbital Rate Dri ve Electro nics for the Apollo Command

Module and Lu nar Module. NASA TN D-7784, 1974.Vernon, J e s s e A. : Apollo Experience Report - Guidance and Control Systems:Lunar Module Mission Pr og ra me r. NASA T N D- 7949, 1975.Shelton, D. H. : Apollo Experience Report - Guidance and Control Systems:

Lunar Module Stabilization and Control System. NASA TN D-8086, 1975.Kurten, Pat M. : Apollo Experience Report - Guidance and Control Systems:Lunar Module Abort Guidance System. NASA TN D-7990, 1975.

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