Apollo Experience Report Guidance and Control Systems Lunar Module Abort Guidance System

8/8/2019 Apollo Experience Report Guidance and Control Systems Lunar Module Abort Guidance System

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C A L N O T E NASA TN D-7990

CASE FILECOPY

EXPERIENCE REPORT -SYSTEMS:

MODULE ABORT GUIDANCE SYSTEMut M e Kzlrten

B. Johnson Space CenterTexas 77058

A L A E R O N A U T IC 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. JULY 1975


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1. Report No. I 2. Government Accession No. I 3. Recipient's Cataloq No.NASA TN D-7990APOLLO EXPERIENCE REPORT

4. Title and Subtitle

GUIDANCE AND CONTROL SYSTEMS:LUNAR MODULE ABORT GUIDANCE SYSTEM

5. Report DateJuly 19756. Performing Organization Code

JSC- 085897. Author(s)Pat M. Kurten

9. Performing Organization Name and AddressLyndon B. Johnson Space Cen terHouston, Texas 77058

13 . Type of Repor t and Period CoveredTechnical Note. Sponsoring Agency Name and Address

8. Performing Organization Report No.JSC S-424

10 . Work Unit No.976- 10-4 1-0 1-72

11. Contract or Grant No.

National Aeronautics and Space AdministrationWashington, D. C. 20546

^. u; n..__19. Security Classif. (of this report)Unclassified Unclassified

ZU. Security Ciassii. io i this pagei 4 1 .73

14. Sponsoring Agency Code

22. Price'$4.25

15. Supplementary Notes

The history of a unique development program that produced an operational fixed guidance systemof ine rtia l quality is presen ted in th is repo rt. Each phase of development, beginning withrequirement definition and concluding with qualification and testing, is addres sed, and develop-mental problems a r e emphasized. Software generation and mission operations a r e described,and specifications fo r the inertia l ref erence unit a r e included, as a r e flight performance results.Significant prog ram observations a r e noted.

7. Key Words (Suggerted by Author(s))- Display Devices- Hardware - Software (Computers)Mission Trajectory t Lunar Landing- Short Circuit Accelerometers.Monte Car lo Method

18. Distribution StatementSTAR Subject Category:12 (Astronautics, General)


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CONTENTS

SectionSUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ACRONYMS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .DISCUSSION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Prel iminary LM Abort Concepts . . . . . . . . . . . . . . . . . . . . . . . .Pr ogr am Redefinition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .System De sc iption . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Test Program . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Software Development . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Abort Guidance System Mission Revisions . . . . . . . . . . . . . . . . . . .Mission Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Abort Senso r Assembly Development P rob lem s . . . . . . . . . . . . . . . .Abort Elec tro nic s Assembly Development Pro blem sData Entry and Display Assembly Development ProblemsAbort Guidance System Software Development ProblemsCapability Estimate . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . .. . . . . . . . . .. . . . . . . . . . .Flight Pe rfo rma nce and Anomalies . . . . . . . . . . . . . . . . . . . . . .

PROGRAM OBSERVATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . .Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Thermal-Va cuum Acceptance . . . . . . . . . . . . . . . . . . . . . . . . .Compatibility T esti ng . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Dual Source Proc urem ent . . . . . . . . . . . . . . . . . . . . . . . . . . .A sst:iilbiy EeiiauiXy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Page

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SectionTransient Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Software Development . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Memory Functions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Scale. actor Asymmetry . . . . . . . . . . . . . . . . . . . . . . . . . . . .Electroluminescent Display . . . . . . . . . . . . . . . . . . . . . . . . . .Pushbuttons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Pushbutton Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Environmental Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Ground-Support -Equipment H ea te rs . . . . . . . . . . . . . . . . . . . . . .Abort Sensor Assembly Mission Acceptance . . . . . . . . . . . . . . . . . .Abort Sensor Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Gyro Fluid . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Gyro Resonant Frequency . . . . . . . . . . . . . . . . . . . . . . . . . . .Engineering Mode 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Assembly Cabling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Vibration Trans missib ility . . . . . . . . . . . . . . . . . . . . . . . . . . .Split-Pin W i r e Wrapping . . . . . . . . . . . . . . . . . . . . . . . . . . . .Progra m Incentives . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Capability Re ention . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Subcontractor Te st Participation . . . . . . . . . . . . . . . . . . . . . . . .Program Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Earth-Orbit Software Programs . . . . . . . . . . . . . . . . . . . . . . . .Test Programs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Mission P erform ance Analysis . . . . . . . . . . . . . . . . . . . . . . . . .Gyro Rundown Time . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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SectionProgr am Evaluation Reporting Technique . . . . . . . . . . . . . . . . . . .Ine rti al Package Alinement . . . . . . . . . . . . . . . . . . . . . . . . . . .Backup Guidance Definition . . . . . . . . . . . . . . . . . . . . . . . . . . .Nailhead Bond Capacitors . . . . . . . . . . . . . . . . . . . . . . . . . . . .Flight Connector Integrity . . . . . . . . . . . . . . . . . . . . . . . . . . . .Work Packages . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Elapsed-Time Indicato rs and "G" BallsCheckout Meetings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Computer Startup Sequence . . . . . . . . . . . . . . . . . . . . . . . . . . .Operator Er r or Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

CONCLUDINGREMARKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . .

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TABLES

TableI ABORT SENSOR ASSEMBLY SPECIFICATION VALUES . . . . . . . . .

I1 ABORT SENSOR ASSEMBLY TIME-STABILITY LIMITS * * *III ABORT SENSOR ASSEMBLY REPEATABILITY LIMITS . . . . . . . . .IV EARTH PRELAUNCH CALIBRATION TIME-STABILITY LIMITS . . . .V DESIGN VERIFICATION TESTING ENVIRONMENTSF O R T H E A G S . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

VI ACCEPTANCE TESTS PERFORMED ON THE AGS . . . . . . . . . . .VI1 ABORT GUIDANCE SYSTEM ERROR MODEL ICAPABILITY ESTIMATE . . . . . . . . . . . . . . . . . . . . . . . .

VIII ABORT GUIDANCE SYSTEM ERROR MODEL I1CAPABILITY ESTIMATE . . . . . . . . . . . . . . . . . . . . . . . .M ABORT GUIDANCE SYSTEM ERROR MODEL IIICAPABILITY ESTIMATE . . . . . . . . . . . . . . . . . . . . . . . .X ABORT GUIDANCE SYSTEM ERROR MODEL IVCAPABILITY ESTIMATE . . . . . . . . . . . . . . . . . . . . . . . .

XI FINAL PIC, FINAL EP C, AND IFC DATA OF THE AGS . . . . . . . .XI1 IN-FLIGHT DETERMINATION OF GYRO AND ACCELEROMETERBIAS IN FREE FLIGHT . . . . . . . . . . . . . . . . . . . . . . . . .XI11 ALINEMENT ACCURACIES OF THE PGNCS DURING THEAPOLLO 9 MISSION. . . . . . . . . . . . . . . . . . . . . . . . . . .XIV A COMPARISON OF THE VG MAGNITUDES FOR ALL BURNS . . . .

Page10111 21 3

1 51 8

46

47

48

495 05 1

5 152

XV A COMPARISON OF THE RENDEZVOUS MANEUVER VELOCITIESAS COMPUTED BY THE AGS AND THE GROUND . . . . . . . . . . . 52XVI ALINEMENT ACCURACIES OF THE PGNCS DURING THEAPOLLO 10 MISSION . . . . . . . . . . . . . . . . . . . . . . . . . . 53

XVII GYRO AND ACCELEROMETER CALIBRATIONS . . . . . . . . . . . . 54XVIII A COMPARISON OF AGS AND PGNCS BURN RESIDUALS . . . . . . . . 54

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Table PageXIX ALINEMENT ACCURACIES OF THE PGNCS DURING THEAPOLLO 11 MISSION . . . . . . . . . . . . . . . . . . . . . . . . . .XX GYRO-DRIFT PERFORMANCE DATA . . . . . . . . . . . . . . . . . .

XXI ACCELEROMETER BIAS IN FREE FLIGHT . . . . . . . . . . . . . . .XXII A COMPARISON OF BURN RESIDUALS . . . . . . . . . . . . . . . . .

FIGURES

Figure1 Abort guidance system components

(a) Data entry and display assembly . . . . . . . . . . . . . . . . . .(b) Abort sensor assembly . . . . . . . . . . . . . . . . . . . . . . .(c) Abort electronics assembly . . . . . . . . . . . . . . . . . . . . .2 Abort sensor assembly asymmetry requirements . . . . . . . . . . .3 Coelliptic rendezvous flight profile . . . . . . . . . . . . . . . . . . .4 Gyro bias history . . . . . . . . . . . . . . . . . . . . . . . . . . . .5 Accelerometer bias history . . . . . . . . . . . . . . . . . . . . . . .

55565 758

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APOLLO EXP ERIENC E REPORTGU DANCE AND CONTROL SYSTEMS:

LUNAR MODULE ABORT GUIDANCE SYSTE MB y P a t M. K u r t e nLyndon B. Johnson Space Ce n t e r

S U M M A R YWith re sp ec t to the lunar module abor t guidance syst em, the Apollo Pro gr amexper ience included the full range of p rogra m development fr om requ iremen ts throughdesign, development, and production and culminated with success fu l flight. A s are su lt of preli minary lunar-abort concepts, a strapped-down ine rti al- sen sor packagefor attitude reference and an open-loop programer were selected to provide asce ntsteer ing for lunar-orbit insertion. A clear pericynthion orbit could be obtained bypostin sertion burns, and rendezvous wa s to be accomplished by means of ex te rnalinformation. A program redefinition in 1964 added the r equire ment f or a clear peri-cynthion orbit from the initial abort burn and the requirement for a rendezvous withinthe lunar module fuel budget without information from so ur ce s outside the lunarmodule. The require d accuracy of the strapped-down package wa s therefore inc reas ed,

and a gene ral -pu rpos e digital computer containing 4096 words of memory re placed theopen-loop programer. A display and keyboard device was added for crew communica -tions with the computer.INTRODUCTION

The lunar module (LM) abort guidance system (AGS) re pr es en ts the first opera-tional usage of a strapped-down guidance syst em . The AGS wa s designed and developedspecifically fo r application in the Apollo LM. The AGS is composed of a strapped-down(or fixed) inerti al- sen sor package, the abor t sensor asse mbly (ASA); a general-purposedigita l computer with speciali zed input/output, the abor t elec troni cs ass embly (AEA);and a display and keyboard device for c re w communication with the computer, the da taen tr y and displ ay assembly (DEDA). The AGS fo rm s the digita l portion of a hybriddigital/analog guidance and cont rol system and is configured t o provide automatic con-tr ol for mission abort resulting from a primary guidance system (PGS) malfunction.

The AGS was supplied by the LM prime contractor under NASA cognizance at theNASA Lyndon B. Johnson Space Cente r (JSC) (formerly the Manned Spacecraft Cen ter(MSC)) . The E ~ C ~ P I I ?espnnsihilit-y nf the AGS was subcont rac ted . The ASA wa s


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fu rt he r subcontracted by the subcont ractor to the ASA vendor. The softwa re fo r theA E A was contracted by MSC directly to the subcontractor.A s an aid to the rea der , where nece ssar y the original units of mea sure have beenconverted to the equivalent value in the Systzme International d'Unit6s (SI). The SIunits a re written fir st, and the original units are written parenthetically thereaf ter.

ACRONYMSA C EAEAAGSASAB P AC& WCARRCDHC D UC E SCMcSICSMD E D AD M C PD RD T OD V TE LEM 1

2

accept ance checkout equipmentabort electronics assemblyabort guidance syst emabort sensor assemblyBethpage, New Yorkcaution and warningcustom er acceptance readine ss reviewconstant differenti al heightcoupling d at a unitcontr ol electro nics systemcommand modulecoelliptic sequence initiatecommand and service moduledata ent ry and display assemblydesign mission computer progra mdesign reportdetailed test objectivedesign verification testelectroluminescentelectromagnetic interference


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E PCFACIFE BFMESFRRFSGSEHSSCHVSUtIAICsIFCJSCKSCLMLOSLSCMSCMSOBOAPERTPICPGNCSPGSPT SA

Ea rt h prelaunch calibrationf i rst article configuration inspec ionfunctional ele ctroni c blockf u l l missio n eng ineering simulationflight readiness reviewflight simulato rground - support equipmentHamilton Standard System Cent erH-vector spin input rectif icationinput axisinterpretive computer simulationin-flight calibrationLyndon B. Johnson Space CenterJohn F. Kennedy Space Centerlunar moduleline of sightlunar surface calibrationManned Spacecraft CenterManned Spacecraft Operations Buildingoutput axisprogram evaluation reporting techniquepreinstallation calibrationpr im ar y guidance, navigation, and control systemprimary guidance systempulse torquing serv oampl ifier

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RCS reaction control systemRTV room-t emperat ure vulcanizingSA spin axisSI Syst5me International d'Unit6sTPI terminal phase initiate

D I SCUSS IO NP r e l i m i n a r y LM A bo r t C onc ep ts

The prel imi nary LM abort concepts included definition of e ar ly r equ ir emen ts andprel imi nary mechanization.Early requirements. - The AGS was developed to provide the capability for safecrew abort fro m the powered descent, lunar su rface, and powered asc ent ph ases of theluna r mission. During these critical mission phases , information cannot be obtainedfro m the command module (CM) o r Ear th that will e ns ure crew safety aft er failu re ofthe prim ary guidance, navigation, and contro l sys tem (PGNCS). The ground rul es thatdefined the prel iminary concepts of the AGS har dware are as follows.1. Crew safety is prim ary. Thi s ground ru le implies that mission abo rt will beinitiated if one additional failure will cause loss of crewmen. On a vehicle having aPGS and an AGS, failure of eit her will re su lt in an ab or t because one additional guid-ance fai lure will not allow safe rendezvous with the CM.2 . Abort capability must exist at all tim es in ca se of a single PGS ailure. Thecapability to in itia te an abort must be independent of the LM phasing with the commandand service module (CSM). The AGS mu st be capable of coping with ini tial condi tionsat any instant duri ng powered d esce nt and asc ent . The capability of launching fr om thelunar surface at any time mu st exist.3 . The AGS must have independent operat ion . The AGS should be designed toope rate without dependence on the PGS dur ing cr it ic al m issi on phases.4. The AGS should be simple. The syst em should be able to ab ort to a clearpericynthion of 12 1 9 2 meters (40 000 feet) at initial injection or at a later correctionif no additional hardware is required for the later cor rec tion . The CSM, through acommunication link to the LM, will be used to provide the LM with information forperforming midcourse corrections . Rendezvous fro m abort should be completed withinthe LM fuel capacity.5. An all-attitude in er ti al re fe re nce should be provided in the AGS. Thisrequirement was established to en sure a reference if the three-gimbaled PGS shouldencounter gimbal lock.

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6. The AGS need not be designed to complete the lunar landing mission. Thisrule d oes not imply that the re is not some point in.time when the LM wi ll land with aPGS failure .7 . The AGS should have the capability of using both the ascent and descent stagesof the LM. Th is provision wa s made to use the fuel in the des cent stage as well as inthe ascent stage after an abort.Modification of th e preced ing ground rules w a s considered to enable the use of asimpl er attitude refere nce system, such as a pair of gy ro s having two deg re es of free-dom. The requ iremen t for the LM to complete the rendezvous with onboard fuel wouldbe modified to have the LM establi sh itself in a clear pericynthion orbit in the plane ofthe CSM orbit. The requirem ent for the LM to abort to a clear orbi t would be modifiedto put the LM on an inte rcep t path with the CSM. Th ese modifications to the groundru le s we re examined and disca rded, pri mar ily because of the i ncrea sed fuel usage withan inaccura te syst em, the increa sed burden on CSM activity, and the unsafe tr aj ec to ri esinvolved.Preliminary mechanization. - The ground rule requiring LM abort to a c learpericynthion orbit and r equir ing onboard fuel to complete rendezvous made an iner tial-quality attitude reference system a necessity. A four-gimbaled inerti al platform anda strapped-down attitude ref ere nce we re considered. The strapped-down ine rtia lrefer ence was selected. The development ri sk s associated with a strapped-downsys tem w er e outweighed by the mechanical simplicity, the inhere nt ruggedness, theease of maintenance, and th e potentia l of the strapped-down sy st em as a lightweight,reliab le sys tem fo r the LM abort mission and for future s pace applications. A t theti me of evaluation, the weight of the strapped-down s en so rs (gyros and accel erom et er s)was est imated at 7 kilog rams (15 pounds); the associated elec tro nic s wer e est imate dat 10.9 kil ogr ams (24 pounds).An open-loop pro gra mer w as selected to provide the asc ent steering neces saryfor a clear lun ar orbit. A pr og ra me r with an estimated weight of 9 kilog rams(20 pounds) wa s the least complex, proven device fo r accomp lishing th is functionbecause AGS navigation wa s not required. The prog ra me r provided a vehicle pitchsequence based on time of abo rt and store d constants and produced a th rust cutoff sig nalbased on the velocity output of an acce ler ome ter mounted along the vehicle thru st axis.A compensation scheme was incorporated into the pr ogram to c or re ct for to leran cesinherent in vehicle propulsion. The compensation scheme comp ared velocity r eadin gsfrom the thrust axis accelerometer taken at specified ti me int erv als with nominalvelocity va lues fo r adjusting pitch attitude and velocity at cutoff. Two ac ce le ro me te rs

mounted nor mal to the th rust axis provided attitude bias ing in co rr ec ti on of vehiclethr ust misalinement. (The system outputs were attitude e rr o r signals for controllingvehic le attitu de, total -atti tude s ign als fo r display, and an engine-off signal. ) Thepro gra mer wa s envisioned as a 2048-word computer having a fixed memory. A controlpanel weighing a pproximately 4.5 kilog rams (10 pounds) completed the preliminarymechanization concept.

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Program RedefinitionThe following prog ram redefinition w as made concerni ng AGS requ ire me nt s andmechanization.Requirements. - In the fall of 1964, a ser i es of LM pro gram redefinition mee tings

was held biweekly to reconfigure the LM guidance and cont rol sy st em s, including theAGS. Several changes we re made to the ground r ul es covering the existing conceptualdesign.1. The AGS should be capable of per forming rendezvous with the CM within theLM fuel budget without information from so ur ce s outside th e LM.2. The AGS should provide information for monitoring the PGS.3 . The AGS should be capable of providing LM ab or t to the CSM fro m any phaseof the LM mission.4. The AGS should be capable of in se rt ing the LM into a cl ea r pericynthionorbi t of 9144 mete rs (30 000 feet) above the Moon. Th is req uir eme nt origin atedbecause the height of some of the luna r mountains was dete rmined to be 8839 mete rs(29 000 feet), although the mountains were not on the lunar equator, the landing site,o r the lunar-orbit location. Th is requir ement also neces sitate d cl ea r pericynthion onthe first abort burn and thus increa sed the accur acy req uirem ent of the ine rtia lreference.5. The AGS attitude ref er en ce should be maintained at 25 deg/sec angularrotation. Thi s req uir eme nt was estab lished because of the vehicle rates necessary forprop er vehicle respon se to manual commands dur ing the final phase of lunar landing.Mechanization. - Revision of the ground ru le s resu lted in discardi ng the open-looppro gramer concept because of inflexibility and incapability. A general-pu rpose com-puter containing 4096 words was selected to rep lace the pr ogr ame r.Navigation capability was added to the software, and an explicit guidance law wasselected to perform the rendezvous. A hard-l ine interf ace between the rendezvousra da r and the computer was added fo r navigation updates. Thi s interface was laterchanged to a manual interface because of a mechanization problem.A connection between the PGS dig ita l te le me tr y downlink and the AGS computerwas added t o obtain state vec tor initialization information f o r navigation; t hi s additionaugmented the existing alinement interfac e for attitude ref ere nce alinement to the PGS.A digital tel eme try downlink wa s added fro m the AEA. Added to the exist ing total -attitude and attitude e rr or displays were computer interf aces for displaying altitude,altitude rate, and lateral velocity for monitoring the PGS.The AGS the refore becam e a full guidance syst em. After alinement fr om the PGSwith attitude information and af te r initi ali zat ion with LM and CSM position, velocity,and epoch time data, the AGS continuously computes LM attitude, LM and CSM position

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and velocity, and abo rt tra je ct ori es fo r predetermined orbital insertion conditions orfor orbital transfe rs to the CSM. The asse mblie s compris ing the AGS are shown infigure 1.

(b) Abort se ns or assembly.

(a) Data entry and display asse mbly.

System DescriptionThe de vic es comprising the AGS arethe AEA, th e DEDA, and the ASA.Abort elect ronics assembly. - TheAEA is a general-purpose, high-speeddigital computer that p erf orm s strapped-down attitude reference computations andnavigation and guidance functions requiredto steer the L M to rendezvous. The AEA

con sis ts of a cor e memory, control andari thmeti c logic, input/output ci rc uit ry ,and a power supply. The ass embly weighs (c) Abort electr onics assembly.14.8 kilograms ( 3 2 . 7 pounds) and hasdimens ions of 6 0 . 3 2 by 2 0 by 1 3 . 3 3 centi-meters ( 2 3 . 7 5 bv 8 bv 5 . 2 5 inches). The

Figure 1. - Abort guidance systemcomponents.AEA (1 ) pr oce ss es input information fro mthe ASA, th e DEDA, and the PGS; (2 ) per-fo rm s attitude refer enc e alinement and state vector initialization; (3 ) calibratesin-flight, iunar sur fac e, and acc eie romete r biases; (4) periorrns b e u - ~ e b ~ b ,3 ) I I I ~ I I -tains att i tude reference ; (6 ) pe rf or ms navigation computations; (7 ) sel ect s mode;

1,- L - - I - . f r \ _ _ _ _ 2. -

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(8) perfo rms guidance computations; (9) outputs st ee rin g and engine comm ands ;(10) proces se s tele met ry; and (11) dis plays outputs fo r in-flight monitoring.Half of the 4096 wor ds of m em or y in th e AEA are hardwired and half areera sab le. Fifty-three esca pe points exist throughout the wired memory so that cor -rections can be made, as required, in softwired memo ry. The memo ry us es

0.0008-meter (30 mil) ferrite co re s wired in 64 by 64 planes, one plane fo r each wordbit. The word length is 18 bits. Data words consist of a sign and 17 magnitude bit s.Instructions con sis t of 18 bit s that contain the o rd er code and a single operand ad dr es sThe memory is a coincident current, parallel, random-access core memory witha cycl e time of 5 microseconds. The functions located in the fixed memor y are thosethat require the fastest computation or those that are mission independent (or bothtypes ); that is, direc tion cosine and tangent rou tin es. Values that might feasi bly varyfrom mission t o mission, such as the guidance equ ations and the ra da r filter, arelocated in the er as ab le memory. Computer computations are done in 20-millisecond,40-millisecond, and 2-second cycles . The direction cosine matri x is updated 50 timespe r second; the attitude e r r o r and engine com mands are computed duri ng the

40-millisecond cyc le; and navigation and guidance are computed during the 2-secondcycle.The comp uter uses a fract ional two's complement paralle l arithme tic section.The computer add time is 10 microseconds; the multiply time, 70 microseconds. TheAEA executes 27 basic instructions. The clock frequency is 1.024 megahertz. A50-word digit al tele me try downlink is outputted from the AEA once per second.Integrated circui ts, thin film networks, and multilayer cir cui t boa rds are usedextensively to minimi ze si ze and weight. A l l sub ass emb lies except the power supplya r e packaged in groups of multi layer board s interconnected by a wiring matri x. Split-pin wire wrapping is used to connect the multilayer board s to the matrix. The matrixca rr ie s all signal lines between the core memory, the arithm etic and control logic,and the input/output c ir cu itr y. Power is distributed by laminated bus ba rs encapsu-lated in the wiring matr ix. The power supply module s a r e of cordwood cons truct ion.The s idepla tes of the AEA are attached to cold rails on which the AEA is mounted inthe L M aft equipment bay. The AEA di ss ip at es 81 wa tt s of power.Data entry and display assembly. - The DEDA is used to cont rol AGS modes ofoperation, to in se rt dat a in the AEA manually, and to command the conten ts of a cces -sible AEA memo ry to be displayed on nume ric read outs. Manual contr ol is accom-plished by depres sion of pushbuttons on the DEDA panel in the requ ired sequence. TheDEDA, AEA, and ASA ex te rn al conf igurations are shown in figure 1.The DEDA cons is ts of two main ass em bl ie s: a contr ol panel housing the e lec tro -luminescent (EL) num eri c display and the data- ent ry pushbuttons, and a logic enclosurehousing the dr ive ci rcui ts , the input/output ci rcui ts , the power conditioning ci rcui ts ,and the logic cir cui ts. The logic enclo sure is a hermetically sealed assembly housingnine multilayer ci rcu it boards. The DEDA ci rc ui try is composed of flat pac ks andthin f i l m networks. The DEDA ha s acce ss to 452 of the 4096 memor y locat ions in theAEA. Manual co nt ro l is accomplished through the use of 4 pushbuttons f o r "clea r, ""readout, " "enter, " and "hold, " respectively; in addition, 10 pushbuttons f or thedigits 0 to 9 and 2 pushbuttons for arithmetic signs are available.

8


16/73

The readout co ns is ts of an EL nine-window display. Th re e dig its are used todispl ay, in oct al fo rm , the add re ss of the memory location into which informati on isto be inserted o r from which information is to be extracted . Six digits, the first beingthe arithmetic sign, are used to display numer ic information. Information tra ns fe rfrom the DEDA to the A E A is accomplished by means of a 36-bit serial digita l word.Display par ame ters fr om the A E A are updated twice per second. The D E D A weighs3.4 kilogr ams (7.5 pounds); the dimensi ons a r e 14.0 by 1 5.2 by 13.18 centi met ers(5.5 by 6. 0 by 5.19 inches ). The operat ing power is 10 wat ts supplied by the A E A .The D E D A is mounted without cold rails in the LM cabin.

Abort se nso r assembly. - The ASA is a strapped-down inertial-sensor unitmounted to the LM stru ctu re and oriented with the coordinate refere nce axes along thevehicle axes. The ASA sens es and mea sur es accelerations along and angular rotationsabout the LM axes, converts these motions to discrete increm ents, and tran smi tsthese inc rem ent s in the for m of p ulse s to the AE A for processing. The ASA consistsof th re e floated, pulse-rebalanced, single-degree-of -freedom, rate -int egrati ng gyros;th re e pendulous, fluid-damped ac ce ler om ete rs and associ ated pulse torquing electron -ics; a frequency countdown subassem bly; coa rse and fine tem per atu re contro ller units;a power supply; and interface elect ronics .The ASA us es cu rre nt pulse torquing to rebalance the output from the sen sors .Vehicle motions detected by the s en so rs ar e converted to altern ating-c urrent voltageshaving a magnitude and a phase proportional to the sense d motion. The output of eachgyro is applied to a pulse torquing servo ampli fier ( P T S A ) that quantizes the signa lsand provides output pulses to the AEA at 64 000 pulses/sec and to the sen sor for torquerebalancing at 1 kilohertz.The gyr os used in the ASA we re specifically developed fo r the strapped-down use .The accele romete rs used are 2401 accelerom eters that replaced the Bell VII units

becaus e of contamination experienced with those units. The se ns or s and th ei r asso -ciated electronics are mounted orthogonally on a beryllium block for temp eratur e con-t rol .two devices: a fast warmu p control to provide minimum ti mel ag to full operationalcapability and a fine tempe ratu re contro l to maintain the critical operating temperatur eof 322 K (120' F). Two sensor s, placed mechanically in diagonal legs of a bridge, areused t o detect tempera ture variations.

Operating temperatur e is maintained by single-point cont ro l and is provided by

Both gyro s and acc ele rom eter s operate on a torque-balance principle and usecommon pulse torque servo amplif iers. Time-modulated cu rre nt is used to offset inputangular rates and a cce ler atio ns by a forced limit-cycle pulse torquing system . Binarytorquing using alte rna te positive- and negative-current peri ods ra th er than dis cret epulses provides a high angular information ra te and a low s ys tem switching rate.Pulses f rom the pulse torque servoa mplifi ers to the A EA are quantized at 2-16 rad/pulse for the g yro s and at 0.000952 m/sec/pulse (0.003125 ft/sec/pulse) for the accel-erometer s and are supplied at a maximum ra te of 64 000 pulses/sec.

9


17/73

The nec es sa ry mounting accu rac y of the ASA is obtained mechanically within2 arc-minutes through four machined feet mounted to the L M navigation base outsideand above the LM cabin. A side-mounted coldplate is used in the ASA for heatremoval. The ASA weighs 9 . 4 kilograms (2 0 . 7 pounds) and has dimensions of12.9 by 2 2 . 8 by 29.2 centimeters (5.1 by 9.0 by 11.5 inches). The ASA power dissipa-tion is 74 watts.The ASA specification values fo r inerti al pe rform ance used in the development ofthe ASA are given in table I. (See fig. 2 also.) The ASA specification criteria are aug-mented by para met er stability c ri te ri a that are exclusively fo r bench-test m aintenance-equipment calibrations at one facility.repeatability limit on the thr ee re ading s used to dete rmine a mean representing theparam eter value for any one calibration ru n is also required fo r a bench calibration.These limits are given in table 111.

The test limits are given in table 11. A

TABLE I. - ABORT SENSOR ASSEMBLY SPECIFICATION VALUES

Gyro scale-factor errorGyro b t u

Gyro lnput ax18 (IA)mass

Gyro spin axla (SamaesUnbala"Ce

UnbahnCeGyro Y-axis alinementAccelerometer Y-axis allnementGyro Z - ax i s alinementAccelerometer Z-urls ahernen1Gyro X-axla alinementAccelerometer X-axis allnementGyro Z-ax18 a l inementAccelerometer Z-aris llnementGyro X-axis ahnementAccelerometer X-axis allnementGyro Y-axis allnementAccelerometer Y - w s allnementAccelerometer scale-f actor err01Accelerometer bias discrepancy

Accelerometer bias discrepancyGyro blaa discrepancylnput vcrtlcal minusoutput verticalspln vartlcal mlnusoutput vertlcalOutput up minusoutput dawn

scale-f ador nonllnearlty

Accelerometer acale-factorwallnearltyGyro scale-factor asymmetryAccelerometer acale-factorulgmmdry

Charnel Trim value

500 pulses/mln2. 4 deg/hr2 . 4 deg/hr2 . 4 deg/hr4 deg/hr/g

4 deg/hr/g

130 are-sec250 arc-see25 0 arc-see250 arc-sec130 Z ~ E - B I C130 arc-see250 arc-sec250 arc-8ec130 arc-8ec130 are-sec130 arc-se c250 arc-sec

L500 pulses/mlnmoug

10Qllg_.

. 2 deg/hr

. 2 deg/hr. 3 5 d e g h r

0 to 10 d e d s e cIO to 22 deglsec22 to 25 deg/sec1wvs

(32%0

Trim stablllty

10


18/73

wCL- 20 0

Because of th e fixed nat ur e of astrapped-down guidance system, completecalibrations requi ring rotatio n cannot beperformed in a vehicle. However, an Ea rt hprelaunch calibration (EPC) progr am usedin the AEA provid es a determina tion of gyrodrift when vehicle azimuth and latitude areknown. In addition to the specif icationrequirement, the time- stability values oftable I V were generated on the basis of theAGS estimated capability in flagging non-

150% pulseslmin

450I , I , , , I , characterist ic E P C results.0 1 2 3 4 5 6 7 8 9 1 0Inpu t rate, deglsec

Figure 2. - Abort s ensor assem bly asym-metry requirements.TABLE n. - ABORT SENSOR ASSEMBLY TlME-STABILITY LIMITS

Parameter

Crro bh s

Gyro scale factor(2.62 eg/see)

Gyro s p i n mass unbalanceGyro input mass unbalance

x gyroc

Accelerometer bias. IA

Accelerometer scal e factor

Accelerometer misalinemei

CtlaMel symbol unit0 to 2

0.35

I60

. 36

. 30- _ 0

. 2 5

46

-w46-

2 o 6

0.48

I95

.44

35-.48.3 5

46

e140

e- 110

46-

i o IO0. 4

240

. 5:

.4!- . 5'.4!

46

e155

-137

46___

Limit on delta-mean fortime mt erva l . days

10 to 20

0.60

300

. 6 6

.64-.66.64

46

e170

e80e-192

46

!O to 40

0 . 6 9

300

. 6 8

.90-1.28

.90

46

195

80,280

46

'H88C = Hamil ton Standard System Center.bBPA = Bethpage, New York (location of Crumrnan Aerospace Corporation),'Applies only dter BO days from seMor acceptance,dThe llmit of 46 can be increased to LB O d an output-axis (OA) apparent misalinement problem has been demonstrated to exIS1esmaller t h a n acceptance test Limits.

40 o 60

0.66

300

1. 03

1. IO- 1 . 6 3

1 10

46

21 5

80352

46

60 o 90

0.69

300

1 . 1 9

1.35-2.011. 35

46

240

80-445

46

0 to 120

0. 3

300

1. 30

1.56-2.391. 56

46

264

BO

-506

46

11


19/73

TABLE III. - ABORT SENSOR ASSEMBLY REPEATABILITY LIMITS~

ASA parameter

Gyro bias

Gyro scale factor(2.62 deg/sec)Gyro IA mass

unbalanceGyro SA massunbalanceGyro IAmisalinement

Accelerometerbias

Accelerometerscal e factorAccelerometer IAmisalinement

Symbol

G G GBiI BiO BiS

SG

U;

SY..1J

ABiIABiO

BiPSA

A

CY..1J

Unit

deg/hr

puls es/min

deg/hr/g

d eg/hr/g

arc-sec

wg

IJ.g

Pg

pulses/min

arc-sec

Limit ondelta -m e as ur em entswithin a set of3 calibrations

0. 29

1 1 2

. 22

.30a25

4 2

76

76

22

20

Limit on standardleviation u froma set of3 calibrations

0. 17

66

.13

.1815a

25

45

45

13

1 2

aThe limit on u for gyro misalinements y x z , yyx, and y can be increasedzxto 106 arc- sec and the limit on delta- measur ements to 180 ar c-s ec for a demonstratedOA apparent misalinem ent problem.

1 2


20/73

0Q,00W0W00*0*00N0cu0

c,-Y-Y-Y0,z-0WY-W0NY-N00c,

0

O Na O Na O Na O Na

~ h

uPIwPIw&

d30cdk

cd0E0cdcdmEQ,

.rlY

.-(

.r(

dc.,

.rl

f$II a

1 3


21/73

Test ProgramThe test progra m for the AGS consi sted of design feasibil ity and verifica tiontesting, qualification testing, spacec raf t testing, full mis sion engineering simulation(FMES), certification te st requir ement s, and acceptance testing.Design feasibility and ver ificat ion te st s. - Design feasibility tes ts w ere incor-porated into the AGS program fo r component and par t se lection; fo r investigation of theperfo rmance of breadboa rd models, components, and suba ssembl ies under va ri ousenvironmental conditions; for select ion of ma te ri als ; and for substantiation of safetymar gin s and other analytical assumptions. The feasibility te st s were conducted exclu-sively by the assem bly vendor in suppor t of design. Design verification te st s wereincluded in the program to substantiate the cor rec tn ess of the assem bly design fo r theintended mission under simulated ground and flight environments and off-designconditions.Four a sse mbl ies we re initially designated as reliability test models and assignedto design verification testing. Because the pro gram schedule and cos t considera tions

necessitated a reduction in hardwa re, only two ass emb lie s were assigned to designverification testing. The first unit was assigned to cr itic al environments, electromag-netic interference (EMI), and ove rs tr es s testing. Cri tic al environments we re to beselect ed by the vendor to en su re that the asse mbl y would p as s the design-li mit qualifi-cation test. During th is test , the environments wer e to be increased from missi onlevels to twice mission levels. The ov ers tre ss test was to be performed until failureand was to use launch o r missi on environments exclusively. Failure -mode predictionanalysis was to be used in selecting the ove rs tr es s environments. The second designverification t es t (DVT) unit wa s assigned exclusively to mis sion simulation in the o rd erin which the e nvironment s were t o be experienced.Progra m constrain ts, pri mar ily schedule problems, result ed in cutting the DVT

progr am to one asse mbly and rest rict ing the te st s to design-limit environmental levels .Preproduction units wer e used f or design verification testing, although the prog ramgoal wa s to use production har dware. The DVT environm ents fo r the AGS a re givenin table V.In addition to design verification testing, a series of design-proof te st s was pe r-formed on the AGS to evaluate i nterfa ces, angula r vibration, gyro scale-factorasymmetry, EPC (with and without sway), navigation in static and dynamic environ-ments, in-flight calibration (IFC), esonance (3 hertz) , coning, constants polarity,and flight-configuration compatibility (test set compared to flight-configurationperformance ).

1 4


22/73

TABLE V.- DESIGN VERIFICATION TESTING ENVIRONMENTS FOR THE AGS

Environment

Therm al vacuumResonance sea rc hTempe rature vibrationShockTemp eratur e/humidityAccelerationLeakCor ros ive contaminantOxygenVibrationGyro and accelerometerscale-factor linearityGyro and accelerometermechanical frequencyresponseAngular vibrationFrequency resp onseSaturation and bottomingCalibration transient sMagnetic fieldsusceptibility

4bort electronicsassemblyXXXXXX

Data entry andiisplay assemb lyXXXXXXXXX

4bort sens orassemblyX

XX

X

XXXXX

15


23/73

Qualification. - Two production AG S units wer e assigned to qualification testing.The qualification pro gram wa s originally divided into two phases , qualification andpostqualification, with a teardown and inspection between the tests.The first qualification unit was assigned to simulate the mission environment.Ground environment tests were to be performed first, followed by two flight environ-

ment simulations . The postqualification te st ing was to cons is t of two additional flightsimulations. The second unit wa s assigned to desig n-lim it test ing and unique environ-ments, such as salt spr ay. The postqualification test fo r thi s unit was t o be twoovers t ress tests.A ground rule of the test program was that the entire D V T program be completedbe fo re the beginning of qualification tes ting.desig n verification test ing and qualification te sting overl apped; thi s ru le wa s modifiedfor the AGS so that an environment was tested in design verification testin g before thesa me environment w as demo nstrated in qualification testing.

Because of schedule considera tions,

Pro gra m cos t and schedule consideratio ns res ult ed in the eliminati on of postquali-fication testing. The number of flight simu lations wa s al so reduced , and the ex po su re swe re lengthened accordingly because of the ti me fo r se tup and teardown of a tes t andthe potential damage resulting from the activity.

In the endurance qualification, a 1000-hour burn-in before start ing the environ-ment test s was established to simulate the a ssemb ly test and checkout perio d. Becauseof schedule con str ain ts, th is period was reduced to 250 hou rs, and E M 1 testing wasperformed during the period.The AGS qualification hi sto ry c onsis ted of the following se par ate tes ts .1. ASA origi nal qualification2. ASA acc ele rom ete r change qualification3. D E D A origi nal qualification4 . D E D A modified E L displ ay mounting qualification5. AEA origi nal qualification6 . A EA modified internal clock qualification7 . A E A incre ased vibration level s qualification8 . A EA modified me mory read/write clock qualification

The multiple qualification tests we re perf ormed because of design changes as noted.Qualification testing was monitored by periodic test revie ws a t the vendor facility duringthe testing.

1 6


24/73

Spacecraft testing. - The subsystem checkout at the prime contractor and at theNASA John F. Kennedy Space Cente r (KSC) was controlled through a test and checkoutrequir ement s document fo r each spacecraft. The test criteria for each vehicle we resupplied to KSC by means of a test and checkout specifica tion and cr i te r ia document.The KSC respon se to the MSC test requireme nts document, outlining the deta il s of thet es t s t o be per formed , wa s stated in the te st and checkout plan. Opera tiona l checkoutproc edur es controlled the detailed p rocedu res for each test.Full missi on engineering simulation. - The prim e c ontra ctor's FMES wa s used toint egr ate AGS hardwa re and sof tware into a closed-loop, six-degree-of -freedomsimulation with other flight control ha rdware and software. The ASA was mounted ona flight-attitude table, and all hardware functions were used except the acce lerom eteroutputs, which we re nec ess ari ly simulat ed fo r input to the AEA. An extensive testprogra m w as performed on the AGS software baseline, the design mi ssion c ompute rprogram (DMCP), followed by tests of eac h flight prog ram with har dware. Gyro ma s sunbalance proved to be a significant e r r o r source that could not be eliminate d in thedynamic tests without complex compensat ions that wer e not re adily available.

activation of the FMES fac ili ty, a three-degree-of-freedom test was performed on apreproduction unit to veri fy the st abili ty of the attitude-control loop.Before

Certification test requirements. - A se ri es of performance demonstrat ions wasset up in the pro gram as con str ain ts on flight. These tests required formal approvalof the test plans, o r certification test requir emen ts, and form al approval of the docu-mented test resul ts, or certification test endorsements. The AGS certi fica tionre qu ir eme nt s included qualification test ing of the ASA, the AEA, and the DEDA, anddemons tra tion of the AGS EPC, in-flight gyro calibration, mission performance fo rthe LM-3 Earth- orbit miss ion, mission performance for the DMCP software pro gra m,lunar mission perfo rmanc e, and in-flight ac cel ero met er calibration. To avoid redun-dant test ing, functions normall y evaluated as part of vehicle checkout were notperformed as certification test requirements.Acceptance tests. - One majo r change in acce ptance testi ng wa s made when thevibration leve ls for e ach operating assembly were changed from 3g root- mean- squar erandom vibration to 6.6g root mean squar e. The change was made to ens ure adequateworkmanship ra the r than to refle ct the wors t mission environment (approximately 3groot mean square ). The acceptance tests performed on the ass embli es are given intable VI. Additionally, as a pa rt of acceptance f or ini tial del ive rie s that were accepte das a system, a compatibility test was performed using the three ass emb lie s that com-pr is e the AGS. The tests perfo rmed on the AGS consis ted of thr ee sets of calibra-tions: a lunar aline, an attitude ref ere nce exerc ise, and a connector waveform test.

17


25/73

TABLE VI. - ACCEPTANCE TESTS PERFORMEDON THE AGS

Tes t

VibrationIThermal vacuumDetailed operation alHardwired memor yWaveform and power supply regulationOperationalSix-position pushbutton activationLow-level vibration and pushbutton

activationTemperatureLeakaLuminescence and pushbutton fo rc eWarmupTes t connector on/off and calibrati onGyro scale-factor linearity andasymmetryVibration (launch and boost) andcalibrationServomotor frequency respon seGyro runup and calibrat ionVibration (ascent/desc ent) andcalibrationPulse modingGyro and accelerometer s cale-factorlinearity and as ymmet ry and calibratioAngular vibrationAccelerometer pendulum stati c frict iontest and calibrationGyro scale-factor linearity andasymmetryCalibration

Lbo rt el ect ron ic sassemblyData entry andisplay assembly

X

XXX

XXX

4bor t s e nsorassembly

X

XXX

XXXXXX

XXX

Xa

1 to 5 hours.To detect leaking pushbuttons, this es t was performed in a vacuum and expanded from

1 8


26/73

SofW.are D eve IopmenSoftware development included the formulation of ba si c req uirement s and devel-opment procedures , verificat ion methods, and use of the softw are .Software functional descrip tion. - The following basic miss ion re qui rem ent sgoverned softw are development.1. Orbital insertion shall be achieved within a pericynthion gr ea te r than9144 meters (30 000 feet).2. Abort to rendezvous fr om any point in the LM mis sio n sha ll be accomplishedwithin the LM fue l budget. The total asce nt differential velocity (delta-V) is1946 m/sec (6386 ft/sec), sub tra cting 7.6 m/sec (25 ft/sec) fo r docking.3. The maximum navigation e r r o r incurred during powered descent , hove r, andpowered ascent shall not exceed 762 meters (2500 feet) and 1.2 m/sec ( 4 ft/sec).4. The AGS must bring the LM within a 9.3-kilometer (5 nautical mile) sphereof the CSM with a navigation velocity e r r o r of l es s than 9 m/sec (30 ft/sec).The fuel constr aint proved to be variable because as vehicle weight increased inthe cour se of the progra m, the available delta-V dec rea sed . The delta-V available forrendezvous after insertion, the prime consideration, decr ease d from 155 m/sec(509 ft/ sec ) to 106 m/se c (349 ft/sec) during the pro gra m. To satisfy the softwarerequi reme nts, the following capabilities we re necessa ry.1. Navigation of the LM and CSM vehicl es2 . Initializa tion of the LM and CSM state vect or s3. Alinement of the ine rti al refe renc e to a selecte d coordinate syste m4. Maintenance of vehicle attitude information with re sp ec t to iner ti al space5. Steering of the vehicle6. Cal ibrati on and compensation of gyro and accelerometer para meter s7. Solution of the guidance equation8. Generation of monitoring data fo r displays and tele met ry9. Automatic in-flight chec k of compu ter memory and logic

To satisf y the se requir ements, the functions of navigation, alinement, calibration,ra da r data proce ssing , guidance routines, and attitude and engine control wereimplemented.

19


27/73

Navigation is perfo rmed in the AGS ine rtia l referen ce fr am e with the origin atthe cen ter of the attracti ng body. Sensed velocity inc reme nt s fro m the ASA are t rans-for med from body to iner tia l coord ina tes and used in the navigation computations whena threshold is exceeded. The threshold preve nts the integration of ac cel ero met er bia sduring coasting flight. The velocity changes caused by gravity are computed, assuminga spherical gravity model for the attracting body. State vector initialization informationon the L M and the CSM is obtained from a hardwire interface with the PGS telemetrydownlink. If the PGNCS is not operational, the data are inputted through the DEDA.Th e absolute time of th e AGS is initialized through the DEDA. A t ime bias existsbetween the AGS and the miss ion tim e becau se of AGS ti me -re gi st er limit ations. Ini-tially , LM tim e was to be based on manning in lunar orbit; a change to mission ti meinitialized at Ear th lift-off requi red a fixed AGS time bias. The t ime bias is kept inthe PGS, and epoch time fo r AGS initialization h as th is bia s subtrac ted before beingplaced on the primary guidance downlink.

Alinement co ns is ts of updating the AEA directi on cosine ma tr ix by one of threemethods provided. The nor mal alinement fra me is the landing site local vertical. TheX-axis pas ses through the landing site to the c ent er of the Moon, the Z-axis is definedby the vector product of the X-axis and the CSM angula r momentum vector , and theY-axis completes the orthogonal tr iad. The PGNCS prov ides continuous attitude infor-mation to integ rator r eg is te rs in the AEA by mean s of hardw ired connections fromcoupling data units fo r the thre e pri ma ry sy stem gimbal angles. Although the AGSmaintains continuous kno&ledge of the pr im ar y guidance alinement, the AEA directioncosine matrix is updated with the information only on DEDA command. A secondalinement capability is that of l unar aline, in which ASA ac ce le ro me te rs are used todetermine local vertical with respe ct to vehicle coord inates on the lunar surfac e.Azimuth information is supplied from a quantity stored at touchdown or f rom a DEDAinput. An azimuth cor rec tio n fac tor can be inputted to cor re ct fo r lunar rotation duringthe stay t ime o r to correct fo r a CSM plane change. The third al inement mode is body-axis aline in which the dire ction cosine mat rix is set to the body refere nce fra me.

Thr ee calibra tion options a r e available in the AGS softw are. The gyro andaccel erome ter calibration option is designed fo r orbita l operations. The gyros arecalibrated to the PGS over a 5-minute period, and the accelerometer readings in freefall are compensated d uring a 30-second period. During in-flight calibrat ion, jetfir ing s are inhibited for ac cel ero met er calibration, and vehicle rates are held below0.075 deg/sec fo r gyro calibration. The second calibration option is lunar surfacegyr o calibration, which req ui re s 5 minutes. The prog ram initially st or es the existingdirec tion cosine mat rix as a ref ere nce for computing dri ft compensation. The inputangul ar increment is corrected fo r lunar rotat ion rate. In the third option, the acce l-er om ete rs ar e calibrated only in free fall during a 30-second period.Radar data are inputted through the DEDA to re duce the e r r o r in the LM statewith respect to the CSM. The ra da r dat a inputted through the DEDA are range andrange rate after nulling the radar pointing error with respect to the Z - a x i s and storingthe direction cosine. A minimum of six range/range-rate updates at int erv als of4 minutes is required f or an acc urate update.(400 nautical miles), nine updates are required.For ranges as long as 741 kilometers

20


28/73

The fi ve guidance options availab le are as follows.1. The o rbita l insert ion option provides ste ering and engine cutoff fo r inse rtingthe LM nto a desir ed lunar orbit.2. The concen tric sequence initiate option compute s the horizontal burn req uir edto satisfy targeting conditions, the desi red termin al phase initiate (TPI) time, and theTPI line of sight to the CSM for a final orbit transfer to the CSM.3 . The constant d iffer entia l height (CDH) routine comp utes the burn di rec tio n andmagnitude re quire d to make the LM nd CSM orbits coelliptical at a predicted CDHtime.4. The T PI routine calculates the trans fer burn f or placing the LM on a directintercept trajec tory with the CSM based on a fixed transfer time.5 . The external delta-V routine a cc ep ts components of a velocity-to-be-gained

Vtained in a local-horizontal s ystem in real time.

vect or input through the DEDA. The vehicle is pointed along the vector and main-G

Attitude errors are generated by the AEA in accordance with selected controlmodes. During attitude ref ere nce alinement and calibration, the errors are set tozero. In the attitude-hold mode, the attitude e r r o r s cause the vehicle to rem ain at theattitude existing when the mode was entered. In the guidance mode, the attitude e r r o r scaus e the vehicle to s te er in accordance with the guidance equations. In the acqu isi-tion mode, the attitude e r r o r s cause the vehicle Z-a xis t o point toward the CSM ofacilitate radar acquisition.An engine-on command f rom the AGS is issued only after a set of co nst rai nts ha sbeen sat isfied; abor t o r abor t s tage is commanded, a pr es et duration of ullage issensed, the automatic discrete is present, guidance control is set to AGS, and guidancesteering is sel ect ed through the DEDA. During powered descen t, the ullage cons tra intis removed by DEDA command so that, ne ar the lun ar surf ace , ignition of an enginemay be insta ntane ous in the event of a flameout.Disp lays of total attitude, attitude er ro rs , altitude, altitude ra te , and lateralvelocity are computed and outputted fr om the AEA fo r cre w monitoring. The AGS

outputs a digital telem etry list of 50 wo rds each second fo r the monitoring of AGSstatu s and performance.A ground-support-equipment (GSE) serv ice routi ne for loading the AEA memo ry,a DEDA pro ces sin g routine fo r AEA/DEDA communications, and an in-flight sel f-t estroutine fo r checking the AEA m emor y and logic com pri se the ma jor software functions.Software development, verification, and use. - The software developed fo r theAGS was obtained by MSC hrough a dir ect contract with the subcontractor. Integration

with the ha rdw are p rog ram was obtained by using an AGS perfo rman ce and inter facespecification fo r softw are development maintained by th e p rim e c ontr actor and bymana ger fo r AGS software development.---.---+:-- +I.-+ +I.- -..hnn-+--ntnv hnrrhirrJrn nrnwram manager he nppninted taskA c y u c s u i i f j c i i a b u i ~UUCIUIILA Q ~ C V I alccz ..-- =- -= - ----

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A deliverable progr am, including mission co nstants and ASA ha rdwa re compens a-tion constants, wa s est abl ished fo r each flight fo r the softwired 2048 words of the AGSmemory. The individual flight pr og ra ms were amended versi ons of a DMCP delive redin March 1967. Thi s pr og ra m, although nonflight, wa s used in the co nt ract or 's six-degree-of -freedom simulation faci lit y to verify AGS per for man ce of t he re fe renc emiss ion and to satis fy the AGS syst em respo nsibi lity of the cont ra ct or .The planned software development at progr am initiation was based on a ser ie s offive design re po rt s spaced over 11 months from the software vendor in response to andinterleaved with th ree dat a packages from MSC to the vendor. The five design rep or tsconsisted of the following documents: Design Report (DR) 1 - Preliminary AnalysisRep ort ; DR 2 - Program Specification and Equation Test Plan; DR 3 - Equation Simu-lation Results Summary, Equations Document, Operating P roc edu res , ProgramCheckout Plan, and Pe rfo rma nce Analysis Report; DR 4 - Program Checkout ResultsSummary and Program Verification Test Plan; and DR 5 - Program Verification TestRe sults Summary, Programed Equat ions Document, Perf o rmance Anal ysis ReviewReport, Prog ram Listing, Binary and Symbolic Card Decks, and Pr ogram Tape.The three MSC data packages contained special mission requirements and therefined mission trajec tory consisting of the mission preliminary r efere nce traje ctory ,the reference tra jec tor y, and the operational traje ctory . The AGS software perfo rm-ance and interface specification contained all other information required for softwaredevelopment.Because the number of des ign re po rt s proved unwieldy and excess ively tim econsuming, five re po rt s wer e combined into th ree. The development schedule was thenrevised, as follows, in t e rm s of ti me befo re launch.1, A t 12 .5 months befo re launch, MSC data package 1, containing the preliminary

reference trajectory and the mission requi remen ts, is delivered to the su bcontractor.2 . A t 11 months befor e launch, subco ntrac tor DR 1, containing the programspecification and the pre limin ary AGS mission pe rform ance analy sis, is delivered toMSC .3 . A t 10.5 months before launch, MSC pe rf or ms a cr iti ca l design review ofDR 1.4. A t 8 months before launch, MSC da ta package 3 , containing the missionreference trajectory, is delivered to the subcontractor.5. A t 7 months before launch, subc ontra ctor DR 2 , including the equation testresul ts , a verification test plan, and a preliminary program , is deliv ered t o MSC.6 . A t 6.5 months before launch, MSC pe rf or ms the first ar ti cl e configurationinspection (FACI) of DR 2.7. A t 5 months befo re launch, sub con tracto r DR 3 , consisting of a verifiedflight program and related documentation, is delivered to MSC.

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8. A t 4. 5 months befor e launch, MSC perf orms a customer acceptance readi-ne ss review (CARR) of DR 3.9. A t 3 months before launch, MSC data package 3 , containing the operationalflight traj ectory, is delivered to the subcontractor.

10. A t 2 months before launch, the subcontra ctor pe rf or ms the final perf orma nceanal ysis based on data package 3 , and MSC perfo rms the flight r ea din ess re view (FRR) .11. A t 1. 5 months before launch, the subcontractor deliv ers the final progr amconst ants tape fo r co mpute r loading.A f t e r deli very of the baseline p rogra m, the development proce dure for thebalance of th e flight pr og ra ms was shortened to begin at 9 months before launch andsimplified as follows.1. A t 9 months before launch, the MSC data package, containing mission

requi reme nts and the refer ence trajectory, is delivered to the subcontr actor.2. A t 8 months before launch, the subcontractor interim design re por t, con-taining the program status, the verification test plan, and the perfo rman ce ana lys is testplan, is delivered to MSC.3. A t 7. 5 months befor e launch, MSC per fo rms the FACI review of the interimde sign repor t.4. At 6 months before launch, the s ubcontractor deli ver s prelimin ary flow-char t changes, operating procedures changes, and a final computer program specifica-

tion. The unverified progra m is deliverable at this ti me f or incorporation in the prim econtractor FMES and the MSC hybrid simulation.5. A t 5 months before launch, the subcontractor deli ver s the progr am, includingverification test res ult s, performance analysis res ult s, the programed equations docu-ment, the operating manual, and a design report. The progra m is delivered in theform of a binary deck and listing, a bench-test equipment loading tape, and an acce pt-ance checkout equipment (ACE) tape f or vehicle us e.6. At 4. 5 months befor e launch, MSC perf orms a CARR of the program deliverypackage.7 . At 8 weeks before launch, the MSC operational tra jec tor y, miss ion data ,and c onsta nts ar e supplied to the subcontr actor.8. A t 5 weeks before launch, the subcontractor deliv ers the final flight pr ogr amwith updated consta nts in ACE forma t and the final per for man ce analys is. An FRR is

performed at MSC.subcontractor software review board approves the r elea se.Befo re subcontractor submi ttal of documentation, an int ern al

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The performance analysis performed f or program delivery 5 months beforelaunch consisted of Monte Carl o ru ns using the ASA specification e r r o r model and theASA capability estimate e r r o r model in mission simulations. The missio n simulationca se s selected we re the most stringent and always included abo rt fr om hover. Theperformance analysis performed 5 weeks before launch used an er r o r model generatedf rom data taken on the flight model ASA.Verification of p roper computer loading wa s accomplished in the followingmanner. The flight program card deck delivered at the CARR was tran sposed by thepri me contra ctor to an ACE tape o r the for ma t used t o load the AEA by the ACE com-puter at KSC. The prime contractor ACE tape was transmitted to the subcontractor forindependent verification and then sent to KSC. The fina l AEA dump, af te r loading withthe constants-updated program in ACE for ma t (supplied by the subcontrac tor 5 weeksbefore launch), w a s compared with the prim e -contr actor-generated ACE tape, and thenoncomparing m emo ry locati ons (noncompares) we re printed out. The noncompares,which represen t the updated missio n constants, wer e compa red with a list of expectednoncomparing locations supplied by the subcont rac tor at the time of the FRR. Anadditional verification proced ure for comparing the original subcontractor binary car ddeck with the final m emory dump was a lso followed f or the first flight but wa s discarde dfor later fl ights because of redundancy.The verification p rocedures used in the cou rs e of p rogram development by thesubcontract or consisted of equation testing, pro gram checkout, verifi cation testing,and performance analysis. Scientific or enginee ring simula tions of the equations in aclosed-loop configuration were designed to demonst rat e equation perf orm ance underboth nominal conditions and vehicle, se ns or , or trajecto ry di spersio ns of 3 standarddeviations (3a). Testing of the coded pr ogram was pe rformed by using a bit-by-bitinterpretive computer simulation (ICs). Th is check verified that the program w asimplemented in accordance with the equations.A closed-loop inte rpr eti ve simulation of th e AEA prog ram and the vehicle flightcha ract eris tic s (ICs flight simul ator (FS)) was used in verification testing. Verificationthat the program w a s capable of guiding and controlling the vehicle in all operatingmodes was accomplished. Special IC s dr iv er routin es also were formulated to augmentthis test. A Monte C arlo analy sis pro gram using 600 cycles was performed to det er-mine detailed AGS perfor mance com pared to missio n require ments .Verification a lso was performed at MSC by using a hybrid computer facility withactu al ASA and DEDA hardware for independent te st c a s e s and by using specia l testcases in the subcontractor ICS/FS progra m. Additional testing for each mission wasperformed by the p rime contrac tor with a six-degree-of -freedom computer simulation

incorporating the AGS as well as other har dware components in the LM flight controlsystem.The selection of hardw are compensation constants for the final pro gram tape wa smade after examination of the dat a hi story of the ac tual flight ASA. In all cases, theexamination result ed in the se lection of the fin al bench calibra tion values with theexception of the ac ce le ro me te r scale fact or, which was extrapolated to the time of themission. The predictable effects of magnet aging on the acce ler omet er sc ale fact or

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enabled thi s extrapolation. The gyro values were checked by Ear th prelaunch cal ibr a-tions perfor med a fte r final ASA installation in the vehicle. The EP C uncertainty of0.37 deg/hr and the stabi lit y exhibited by the ASA'led to maintaining the compensationvalues at those obtained by the mor e ac cu ra te bench calib ration. In-flight calibrationof gyr o and acc ele romete r biases before undocking in lunar orbit allowed final compen-sations fo r the mission.Other software deliv erabl es supplied t o MSC we re EP C ta pe s and simulated flightproce dures . Ea rt h prelaunch calibration w a s a 20-minute special program fo r dete r-mination of gyro dri ft in the vehicle sway environment. The other ASA par ame te rsused in the E PC tape fo r compensation wer e derived fr om dat a taken dur ing the firstASA bench calibration at KSC.The s imula ted flight program cons isted of a set of p roce dur es used with a speci-fied flight progra m to simulate various mission phases. The purpose of the s imulatedflight was to gain confidence in AGS opera tion with interfa cing LM subsys tems. Simu-lated flight pro ce dur es of t he AGS we re designed to test AGS simulated flight initializa-

tion; CSM acquisition; radar filter; abort from powered des cent; abort fr om the lunarsur fac e; and coel lipti c sequence initiate (CSI), CDH, TPI , and ex te rnal delta-V guidancesolutions. The criteria for the te st s wer e obtained from interpretive computer simula-tio ns and re sul ted in value bounds o r bounded c urves fo r vari ous AEA and displaypa ra me te rs . Simulated flight s we re performed throughout AGS test ing as an integralpa rt of operat ional checkout.Software changes we re accomplished through the su bmi tta l of softw are changepropos als generated independently o r on request by the software vendor. A fo rm alMSC board, con tro lling both AGS and PGNCS software, reviewed the software changeproposals.

Abort Guidance System Mis s ion Re v is ionsRevis ions to the AGS missio n are discussed in the following paragr aphs. Themajor change consisted of replacement of the direc t-as cent tra jec tor y fo r rendezvouswith the coel lipti c flight plan.Direct ascent. - The first AGS rendezvous concept was a minimum-time, direct-asce nt traje cto ry t o the CSM following abort. The CSM orbital altitude fo r thi s missi onwa s 148 kil ome ter s (80 nautical miles). Fo r CSM phase a ngles that precluded d ir ec tascent, the AGS was placed in a parking orbit at the insertion altitude and then at

15 240 me te rs (50 000 feet) to obtain the maximum catchup rate to the CSM. Data fororbital tra nsf er to the CSM were obtained from ex terna l sources.Coelliptic flight plan. - In January 1966, the direct-ascent concept w a s discardedand repla ced by the coelli ptic flight plan, which es sen tia lly provided fo r an LM parki ngorbi t for all asce nt case s. The CSM orb ita l altitude wa s changed to 111 kilometers(60 nautical miles) for th is procedure. A diagram of the coelliptic seque nce is givenin figure 3 .

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= LM orbit-CSM orbit= t imeof ign i t ion of- l ine-of-sight anglesecond bur n

Figure 3. - Coelliptic rendezvous flightprofile.

The AGS, following abo rt o r ascentfro m the lunar surface, ins er ts the LM intoa lunar orbit based on insertion altitude,insertion altitude rat e, and desired lunarorb it, usually 17 by 56 ki lom et er s (9 by30 nautical miles), under a utomatic control.The LM coasts until the time of ignition ofthe first burn tigA, which places the LMapproximately 90" f rom the inser tion point.At that time, the rendezvous profile is ini-tiat ed with the CSI maneuver. The burn isnorma lly done in the e xte rna l delta-V modeusing the reac tion cont rol sy ste m (RCS), asare all burn s of the r endezvous sequence.The CSI burn magnitude is deter mined by aniterative technique such that a desired rela-tionshi p between th e CSM and the LM ex is ts at a specified time (time of ignition of thethird burn tigc) following a maneuver t o produce a CDH between the CSM and the LM.

The CDH burn following CSI is performed at the predicted t ime of LM orbi t apofocus(o r perifocus). Because of th is burn, the LM orbi t bec omes coell iptic with tha t of theCSM, the line of aps ides of the two or bi ts is alined, and a constant differential altitudebetween the two vehicles ex is ts . The final planned burn of the sequence, TPI, is per-formed at the point at which a desi red line-of-sight angle (usually 26.6") between theLM and the CSM is reached. The magnitude is based on a fixed tra nsfer t ime to ren-dezvous, usually equivalent to a 130" centra l angle transf er. The pa ra me te rs of thetransf er trajectory are determined by iterat ion to bring the LM to the predicte d positionof the CSM at the specified time. Two midcou rse maneuv ers are performed, asrequired, after TPI, and the burns are computed by the sa me technique used f or theTPI maneuver. The braking maneuver f or docking is originated at a range of 9 .3 kilo-me te rs (5 nautical miles). A range and range-r ate chart profile is followed, inter-leaved with nulling of the line-of-s ight rate.

The sequence of e vents following an abo rt depends extensively on the time ofabort. The nominal sequence outlined is used for aborts near the lunar surface. Forab ort s before powered descent, t he TP I maneuver may be performed initially. Thenominal sequence is als o subject to extensive changes in the c our se of missi on planning.A fixed o r canned external delta-V maneuver can be performed before CSI; a dual CSIburn may be perfo rmed before CDH; and phasing maneuvers may be incorporated asrequirements for landing site selec tion, rendezvous lighting, and rendezvous locationdictate.

Mission OperationThe AGS is maintained operational in a ready state throughout the lunar miss ionas a backup to the PGS. A nominal operational profile is described in thi s section.

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Lunar orbit. - The AGS is inoperative during the tran sluna r-coa st phase, and theASA heaters maintain the ASA at a tem per atu re of 322 K ( 120" F). Coolant is not sup-plied to the ASA coldplate . After the crewmen enter the LM in lunar orbit, the s yst emis activated and the gyros are tested for a minimum of 25 minutes. An AEA self-testconsisting of an AEA memor y sum check and logic test is performed by DEDA com-mand. The DEDA operation and the EL segmen t operation ar e verified. The AEAtime is initialized by a DEDA entry, and an LM and CSM state vector initializationfr om the PGS downlink is performed. The AGS is then alined to the PGS, and theresul ts are checked by monitoring the total-attitude indicators and the system digitaldownlink. The init ializat ion of the AGS is checked by calling up the range and therange rate to the CSM and the in-plane angle between the Z-body axis and the localhor izonta l on the DEDA and by compar ing the PGS readings . In-flight ca lib rat ion ofthe AGS gyros and accelerometers is then performed.

Powered desce nt. - The gyro drift and navigation e r r o r s of the AGS are mitigatedfo r the descent phase by a state vector initialization from the prima ry syste m 10 min-utes befor e powered desc ent and by anali neme nt to the pri ma ry syst em 5 minutesbef ore powered descent. During powered descent, the AGS is in the followup attituderefe rence mode because the primary system is in the vehicle-control mode. The AGSindependently maintains an attitude refe ren ce, navigates, and solves the ab ort problem.The attitude, altitude, and altitude ra t es of the AGS are compar ed with those of thepr imar y syste m during descent. The LM inertial velocity is displayed on the DEDA.

A t an alti tude of 1829 me te rs (6000 feet), the AGS navigation e r r o r s are furtherreduced by a DEDA alti tude update. An AGS abort may be accompl ished at any time byswitching guidance con tro l to AGS and pr es si ng the abo rt button.Lunar surface. - After touchdown, the AGS lunar sur face azimuth is stored. Fol-lowing a decision to stay, the LM state vector is initialized using information stored in

the AEA before descent that ref lec ts the luna r surfac e state, and the AGS is placed inlunar aline. When all systems are verified as operational, the AGS is initialized andalined to the primary system, the lunar azimuth is again stored, and an AGS lunarsurface gyro calibration is perfo rmed . The AGS is then placed in standby with onlythe ASA ope ra tional for the duration of the stay on the lunar surface.Before ascent , AGS time , state vectors, and alinement are again initialized, anda lunar sur face gyro calibration is perf orme d. The AGS is placed in lunar alineapproximately 30 minutes before lift-off and is changed to guidance steering 2 minutesbefore lift-off to red uce gyro drift effects on the attitude re fere nce.Powered ascent. - During ascent, the AGS is in followup and is independentlynavigating and solving the insertion problem. The thr ust a xi s VG is monitored on the

DEDA during the latter half of ascent. Initially, AGS to ta l velocity is monitored.After insertion cutoff, the AGS velocity residuals in each axis ar e c ompared with thePGS and tri mm ed .Rendezvous sequence. - After insertion, the AGS is updated in relation to the CSMby the inse rti on of ra da r range and ra nge- ra te information through the DEDA. The

P S S facilitates rendezvous radar acquisition of the CSM by a mode that points the

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vehicle Z - a x i s at the CSM. Nine ran ge and range- rat e points are required before thefirst coelliptic sequence maneuver, CSI; six points are required before each subsequentpowered maneuve r. The AGS independently so lves fo r each powered mane uver andoperates without dependence on the PGS after insertion for rendezvous.

Ab o r t Se n s o r As s e m b l y D e v e l o p me n t P r o b l e m sThe following significant problems were encountered during the development ofthe ASA.Initial ac cel ero met er selection. - The Bell IIIB acce lerom eter was originallyproposed f o r the AGS and approved. At the time of pro gram initiation, however, st rongarguments we re advanced by the c ontra ctor f or switching to the Bell VII, a new,scaled-down ver sion of the Bell IIIB. Although the Bel l IIIB operat ional and test experi-ence was extensive during the Ranger Prog ram and at the NASA Je t Propul sion Labora-tory, the ar gument s fo r sa vings of 1 .1 kilograms ( 2 .5 pounds) in weight, 0.0005 cubicmeter (2 9 cubic inches) in volume, and 4 watts of steady-state power wer e attrac tive.

The deciding argument, however, was that herme ti c s ealing of the Bell IIIB would beexpensive in t e r ms of scheduling, req uir ing 16 weeks fo r delivery of an unsealed unitand 8 months for delivery of a sealed unit c ompared to 4 weeks fo r initial sealedBell VI1 del ive ry. A change w a s made to the Bell VI1 unit. Subsequent events , how-ev er , revealed that the hermet ic sealing of the Bell VI1 had not been adequately proved,and a redesign of the ca se was req uired to prevent "oil canning. " Subsequent biasstab ility and contamination p rob lems with the Bell V II have indicated that the originalselection of the Bell IIIB acc ele romete r was c or re ct .Spin motion rota tion detect ors . - The neces sity for monitoring gy ro wheelspeedbecame apparent in the early stages of the progra m. Platinum cobalt slugs were addedto the gyr o wheels with a pickoff c oi l to gener ate a pulse for each wheel revolution.The spin motion rotation det ect ors al so serve d to check wheel runup and rundown ti me sand, therefore, bear ing integrity.Noise. - A general electronic noise problem within the ASA became apparent withthe integration of the first ASA. Ine rti al per forman ce changed when the test -conne ctoroutputs were capped instead of being connected to the te st se t. No single designchange eliminated noise problems. Corr ect ive actions taken included eliminatingground loops with the test s et connected, rerou ting sens itive internal leads, twist-ing instrument torquer leads instead of using shielded cable, using shie lds betweenthe pulse torque ser voampl ifi ers and the power supply, and connecting secondary powerground internally to ch as si s ground.Tempe ratur e maintenance. - The ASA i nst rum ent s we re heated through a beryl-lium block rat he r than through individual se ns or heate rs. The beryllium block re pr e-sented an extensive design task, req uir ing comple x shaping and dril ling to achieveproper heat tr an sf er and the rma l gradients . Nine desi gns for the beryllium block wer erequir ed to achieve the final design.

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Internal cabling. - Flexible, multilayer printed cir cui t cable wa s used in theinterconnecting inter nal harnes sing of the first preproduction ASA primarily becauseof the weight and volume lim ita tions of conventional wiring. Th is method required mor evolume than anticipated, had an unsati sfa cto ry appearance because the ca ble would notsta y molded around co rn er s, and wa s imp rac tic al fr om the standpoint of accommodatingdesign changes. Conventional point-to-point stranded wiring in bundles wa s used in thesecond preproduction ASA.The rma l design. - The ASA power supply was located between the main housingand the hea t sink. The preproduction power supply w a s potted in a th er ma l conductingcompound that allowed heat to permeate through the compound from the main housingand the power supply to the heat si nk surfa ce. The the rm al conductivity of the pottingcompound, however, wa s not as high as anticipated (0.5 W/m-K (0.3 Btu/hr/ft-OF)compared to 36.5 W/m-K (21 Btu/hr/ft-OF)), and a weight penalty was incurred becauseof the dense potting compound. The production power supply wa s designed such tha t allsignificant heat-producing components and modules wer e heat sunk di rect ly to the cold-plate side of the power supply housing. A less dense potting compound could then beused with cons iderab le weight reduction. Heat from the main housing wa s routed bytwo st uds through the power supply to the coldplate surfac e. These stu ds incorporatedadjus table sl ugs by which the th er ma l path between the main housing and coldplate couldbe var ied . The capability of trimming tot al power consumption after ASA assemblyw a s gained.

Mounting feet. - Ea ch pair of preproduction ASA mounting feet was an integralunit at each end of the housing. Each foot contained a Mycalex rin g that ser ved as athermal barrier, but machining the Mycalex was difficult. Bec ause of continuedexposure to the specified vibration environment, fatigue fa ilures occu rr ed . An individ-ual suspension sys tem was incorporated for the production design. Each foot wasattached to t he block with a the rma l path sma ll enough to eli minate the need f orMycalex. Th e body of each titanium foot w a s hollowed, and an aluminum slug covere dwith an e la st omer was inser ted to provide damping and to prevent exceeding the speci-fication resonance requirements.

A second change was nec ess ary when high reso nan ces occu rred as a res ult ofel as to me r deter iorat ion caused by extended periods of vibration. Sever al other mat eri -als we re t est ed for damping efficie ncy when insert ed in the foot in place of the aluminumslug and the ela stomer . Although ana lys is indicated otherwise , a solid titanium footdid provide adequate damping and was less difficult to manufacture. Solid titanium wa sincorporated into the foot design.Cover support. - The preproduction ASA cover that ser ved as the support membe rfor the t hr ee connectors was undesirable structurally and tim e consuming to remove.A sepa rate str uct ura l member wa s designed to support the connectors, and the coverwas redesigned.Accelerometer scale-factor temperature sensitivity. - The Bell V II accelerometerscale-factor temperature sensitivity w a s higher than toler able . Tem per atu re compen-sation was incorporated into the accelerometer design, but the acc eler omete r was laterdiscarded because of contamination problems.

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Thermal runaway. - A potential th er ma l runaway exist ed in the ASA design if thesensing the rmi sto r controlling the fast war mup opened o r i f the t r im resis tor s hortcircuited. Thi s condition could endanger other equipment on the LM adjacent to theASA. A normally closed therm ostat se t to open at a temper ature of 339 K (150" F)(322 K (120" F) norma l operation) and set t o close at 328 K (130" F) wa s incorporatedin series with the +12-volt ASA power supply caution-and-warning (C&W) signal. TheASA is turned off manually if the display indicates a malfunction on thi s line.Accelerometer warmup. - The Bell VII acce lerom eter bias had an excessivelylong warmup time . The requ ire men t fo r *30pg of steady stat e in 25 min utes wa sviolated by s eve ral hundred millig als (pg's) for se ve ra l hours. The long warmup wa scaused by an electrostatic charge buildup between the capacitor plates and the torquercoil on the pendulum when the torquer coil potential was 16 volts above ground and bya change in the resulting electrostatic forces on the pendulum as the c harge leaked offgradually.The acce lerome ter torquer is operated at a 16-volt, dire ct- cur ren t potentialre la tive to ground because of the PTSA design. Th is potential caus es buildup of an

elec trosta tic char ge between the pendulum and the capacitor plate s that is inverselyproportional to the re lat ive distance between them. When the pendulum is not physi-cally centered, one plate exerts a gre at er f orc e on the pendulum than the o ther.Because of low leakage rates of the bridge circuit capacitors, much time is requiredto null the electro stati c forc es. The long time constant c au se s the effective bias tochange proportionately.Initially, a capacitor adjustment was made to bring the e lectr ical null aroundwhich the acce ler ome ter oper ate s into coincidence with the elec tro sta tic null. Bleedresis tors were later added to drai n the ele ctr ost ati c charge.Cover design. - Prob lems developed in building the single-pie ce ASA cover. Theinitial procedure w a s to build the inner skin first, bond honeycomb co re ma te ri al to it,and then bond the out er skin to the honeycomb. To fac ili tat e manufacturing, the honey-comb core mate ri al was changed to polyurethane foam. The foam, however, separ atedat the bond line during extended vibration, and the c or e ma te ri al was changed back tohoneycomb.Additionally, during environmental testing, pins fro m the elect ronic su bas sem -bl ies shorted, through tol era nce buildup, to the inner aluminum skin of the coverthrough an insulating fibe rg lass board. The inner aluminum skin of the cove r wa sreplaced with fiberg las s. The th er ma l and EM1 eff ect s of the change we re acceptab le.Thermal-vacuum effects. - Because of thermal-vacuum exposure as high as0.7 deg/hr, gyro bi as shif ts occurr ed during initi al ASA test ing.first thought to be ther mal bec ause the te mpe rat ure difference between the gy ros and

the mounting block increased as much as 1.1 K ( 2 " F) under vacuum conditions. Thegyro mounting holes were enlarged to minimize changes in heat flow ac ro ss the a ir ga pfr om the gyro ca se to the mounting block between ambient and vacuum conditions. Th isaction did not co rr ec t the problem. A the rmal ly conductive compound was then appliedto the gyro seat s to prevent the therm al r esi sta nce ac ro ss the interface from beingaffected by pre ss ur e variations. Thi s addition reduced the thermal-vacuum shift sbelow the 0.3-deg/hr specification.

The problem was

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Gyro mass unbalance instability. - Both long- and sho rt -t er m m a s s unbalanceinstabilities became evident in the RI-1139B gyro. The long-term shift s violated theiniti al specification requi rem ent of 2 deg /hr /g; thus, the use of m a s s unbalance bec amean uncertain indicator of bearing wear . The s h i f t s wer e orientation dependent and pr e-dict able with orien tation change. The cause of the shi fts wa s pre sumed to be sett ling ofthe two-cut gy ro flotation fluid under gravity conditions. Gyro circle tests were insti-tuted during acce ptance to dete rmi ne that the shift magnitude was bounded, and thespecification w as relaxed to a value of 4 deg/hr/g. Because of the fixed se ns or s, theASA has the advantage tha t only X-gyro-spin-axis m a s s unbalance (due to thrusting) isa significant er r o r sou rce in the mission. Corrective action wa s taken to sto re boththe ASA units and the gyros be

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Apollo Experience Report Guidance and Control Systems Lunar Module Abort Guidance System

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