A Reproduced CopyOF

Reproduced for NASA

^ the

Scientific and Technical Information Facility

ThiMaaterial contains^Snformation atteei-ing thei^onal defence of the United Stateswithin theratenina^f the espionage laws,Title 18, U.SiCTMections 793 and 794, the

of which in anyn a u l h o i d person is pro-

transmissionmanner to ahibited by

(HASA-CR-60659) EXTENDED APOLLO SISTBHSUTILIZATION STOBY* , VOLpUB 14: ,G0^DA^CE

N A V I G A T I O N (General rioters' Corp.) 134 p& N D

N75-72113

Unclas

SPARK Pll OTORS CORPORATION

Prepared for

Manned Spacecraft CenterNational Aeronautics and Space Administration

Houston, Texas

Contract No. NAS9-3140

Copy No.

Subcontractor Report No.

SID 64-1860-14

i/EXTENDED APOLLO SYSTEMS

UTILIZATION STUDY

Final Report

Volume 14

Guidance and Navigation

16 November 1964

Prepared by

AC Spark Plug

The Electronics Divisionof

General Motors Corporation

Milwaukee, Wisconsin 53201

SID 64-1860-14

This document contains lnformatl<the meaning of the espionage lawa. Tilor revelttiaa of which la any

3 national defence of the United States withinU.S.C. Section* 793 and 794. the transmission

i unauthorised person la prohibited by law.

SPARK PlUG-th. f GENERAL MOTORS CORPORATION

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SPARK PLUG-tht Electronics Division of GENERAL MOTORS CORPORATION

CONTENTS

Section Page

FOREWORD 1PREFACE 3

I INTRODUCTION 11PURPOSE 11SCOPE 11CONCLUSIONS 11

II REQUIREMENTS AND DEFINITION 13REQUIREMENTS ANALYSIS 13

Mission Profiles 13Reliability Requirements 13Design Constraints 13Assumptions 16Environmental Conditions 17

BLOCK II G & N System Definition 18Inertial Subsystem 20Optical Subsystem 26Computer Subsystem 30

MISSION OPERATION 34Launch and Parking Orbit Injection 34Parking Orbit 35Translunar Inject 35Transposition and Docking 35Translunar 36Deboost to Lunar Orbit 37Lunar Orbit Corrections 37Lunar Mapping 37Transearth Injection 37Transearth 37Entry 37

III TECHNICAL ANALYSES 39BLOCK II PERFORMANCE ANALYSIS 39

—, Lunar Orbit Navigation 41] Local Vertical Stabilization 51

Block H Reliability Versus Time 52

H

SID 64-1860-14

SPARK PlUG-th« EUctronlci Division of GENERAL MOTORS CORPORATION

MSection Page

PROBLEM AREAS 55 | |Performance 55Reliability 55 ffi

LIFE EXTENSION ANALYSIS 57 f jAreas of Improvement 57Spares /Redundancy 59 pMaintenance Analyses 63 [ jParametric Studies 74

RECOMMENDED SYSTEMS 80 nChanges From Block I System 80 \ iAreas Requiring Further Engineering Development 80

INTRODUCTION 83Scope 85 TjSummary 85 - L 1

TEST PLANNING 87Design and Development Testing 87 HAcceptance and Delivery Testing 89 ! )Interchangeability Testing 93Qualification Testing 97 H

V DEVELOPMENT PLANNING 111OBJECTIVE 111 HSCOPE 111 i I

Management 112Reliability and Quality Control 112 f]Engineering 112 * • 'Field Operations 113

TEST SCHEDULES AND MILESTONES 113 f|

VI CONCLUSIONS AND RECOMMENDATIONS 117CONCLUSIONS 117 'IRECOMMENDATIONS 117 '

VII APPENDIX 119

SUPPORTING DATA 119

n

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SPARK PlUG-tnt Electronici Division of GENERAL MOTORS CORPORATION

ILLUSTRATIONS

Figure Page

1 System Evolution 42 Apollo X Concepts 73 Apollo X Study Approach 94 Apollo Block II Guidance and Navigation System 195 Inertial Attitude Reference 216 16 PIPA Loop 227 Electronic Coupling Unit 238 ISS Coarse Align Mode 259 ISS Attitude Control and Fine Align Modes 25

10 Optical Subsystem—Zero Optics Mode 2811 Optical Subsystem — Manual Direct Mode with Scanning

Telescope Trunnion Angle Offset 25 Degrees 2912 Optical Subsystem — Manual Resolved Mode with Scanning

Telescope Following StLOS 3113 Optical Subsystem — Tracker Operate — Manual Optical

Subsystem 3214 Maximum Position Uncertainty (la) 4215 Maximum Position Uncertainty (la) Versus Initial Accuracies 4316 Maximum Position Uncertainty (la) Versus Angular Accuracies 4417 Altitude Uncertainty (la) Versus Angular Accuracies 4518 Altitude Uncertainty (la) Versus Initial Accuracies 4619 Maximum Position Uncertainty (la) Versus Ground Tracking

Accuracies 4720 Altitude Uncertainty (la) Versus Ground Tracking Accuracies 4821 Maximum Position Uncertainty (la) 4922 Altitude Uncertainty (la) 5023 Reliability Versus Time Per Profile 1 Apollo X Mission

(Block II G & N System Without Modification or Spares) 5324 Mechanization to Allow Suspension Turn-Off 5825 Mechanization to Allow PIPA Turn-Off 5826 Proposed Redundant Electronic Coupling Unit (ECU) Diagram 6227 Malfunction Isolation Procedures 6828 Reliability Versus Time, Profile 1, Apollo X Mission 7529 Reliability Versus Time, Profile 2, Apollo X Mission 7630 Reliability Versus Time, Profile 3, Apollo X Mission 7731 Reliability Versus Weight, Profile 1, Apollo X Mission 7832 Reliability Versus Weight, Profile 3, Apollo X Mission 79

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SPARK PlUG-tho Electronics Division of GENERAL MOTORS CORPORATION"I ;

Figure Page

33 Apollo X G & N Program Progression 84 j34 Manufacturing Flowgram 91 '35 Interchangeability Test Flowgram 95 - ,.- :36 Apollo X G&N and Module Qualification Flowgram Test Sequence 100 { '-37 Apollo X Spare IMU Functional and Environmental Qualification

Test Flowgram 10138 PSA Functional and Environmental Qualification Test Flowgram 102 j39 ECU Functional and Environmental Qualification Test Flowgram 10340 Apollo X D & C Group Functional and Environmental Qualifi- -^

cation Test Flowgram 104 j41 Apollo X —Major Milestones 11442 Apollo X — Integrated Equipment and Test Schedules 115 -x ;

A. 1 IMU Alignment 124 MA. 2 IMU Alignment —Update 125A. 3 IMU Alignment —Update 125 -]A.4 Thrust Vector Control 126 '»

i

A. 5 Midcourse Correction —Orbit Correction 126A. 6 Landmark Tracking 126 f-jA. 7 Attitude Control 126 !/A. 8 Boost Monitor 126A. 9 Nothing 126 HA. 10 Star-Landmark Measurement 127 ]A. 11 Star-Landmark Measurement 127

11

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- vi - SID 64-1860-14

SPAIK PlUC-th. Electronic* Division of GENERAL MOTORS CORPORATION

•} TABLES

Table Page

1 Apollo X Mapping Mission Profiles 142 Lunar and Earth Polar Orbit Mission Local Vertical Stabilization

Cycle 153 Mission Probability of Success 154 Apollo Block II G & N System Error Sources 405 Local Vertical Stabilization Error (la) 516 Downlink Telemetry Signals 677 Apollo X Qualification Summary 998 G & N Equipment Signal List 107A.I Apollo X — G & N Time Line 120A. 2 Block II Module Failure Rates 123

- vn - SID 64-1860-14

SPARK PLUG-tht Electronics Division of GENERAL MOTORS CORPORATION

FOREWORD

This volume constitutes a portion of the Apollo X Study Final Report (SID 64-1860)prepared as part of the Extended Apollo Systems Utilization Study conducted by theSpace and Information Systems Division of North American Aviation, Inc., for theNational Aeronautics and Space Administration's Manned Spacecraft Center under Con-tract NAS9-3140, dated 6 July 1964. S & ID acknowledges the outstanding technicalcontributions made to the study by a number of companies; these organizations areidentified below along with the title of the report for which they are responsible.

The final report has been prepared in the series of 23 volumes listed below:

1. Summary

2. Mission and Performance Analysis

3. Experimental Programs

4. Configurations, Structures, and Weights

5. Mission Plans and Functions

6. Environmental Control System (AiResearch)

7. Fuel Cells (Pratt & Whitney)

8. Alternate Power Source (General Electric)

9. Cryogenic Storage System (Beech Aircraft)

10. Service Module SPS Engine (Aerojet)

11. Command Module RCS Engine (Rocketdyne)

12. Service Module RCS Engine (Marquardt)

13. RCS Propellant Tanks (Bell)

14. Guidance and Navigation (AC Spark Plug)

15. Alternate Guidance System (Autonetics)

16. Guidance Computer (Raytheon)

17. Stabilization and Control System (Honeywell)

18. Communications and Data System (Collins Radio)

19. Earth Landing System (Northrop Ventura)

20. Subsystems Supplement

21. Reliability and GSE

22. Development Planning

23. Condensed Summary

_l_ SID 64-1860-14

Thle document conUloe Informttton^ttQOtlwtho national dcifonee of the United SUtee withinthe muolnf of the •»plon»ge !«*•. ™§n§i :5' ° 8oot'oni ™ ODd 7M' "" "••"•millionor revelation of which In any mannj^l5«a«an»uthorlioci pcrton U prohibited by !•».

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SPARK PLUG-th* Electron of GENERAL MOTORS CORPORATION

PREFACE

The NASA Manned Spacecraft Center has for the past several years examined theapplication of the Apollo spacecraft to missions alternate to the basic lunar landingmission for which it is currently designed. The overall objective of these studies isan assessment of the advantages and disadvantages associated with the application ofdeveloped hardware to these other potential missions.

NASA studies of near-term applications, which represent possible initial extensionsto current Apollo capabilities, include use of the Apollo spacecraft as (1) a space stationlogistics resupply vehicle (carrying up to six men), (2) an earth-orbital space stationor experimental laboratory, (3) a lunar mapping and survey vehicle, etc. This reportpresents the results of the most recently concluded studies of the Extended-MissionApollo, in which the application to a 45-day earth orbital laboratory role and to anextended lunar orbit mission have been examined in depth.

The initial Extended-Mission Apollo study, initiated in August 1963 under contractto NASA/MSC, examined the suitability of Apollo as an earth-orbital biomedical/behavioral experimental laboratory. These experiments were to provide a basis fromwhich man's suitability for protracted space missions could be determined. Three basicconfiguration concepts, indicated in Figure 1, were investigated throughout thestudy. Configuration Concept I utilizes only the Apollo command and service modules(CSM), with experimental in-orbit work space made available in the command moduleby the elimination of one crewman from the current crew size of three. The ApolloCSM subsystems were modified to sustain orbital operations for periods of up to 120days without resupply. Configuration Concept II consisted of the Apollo CSM plus a5600 cubic foot laboratory module built within the geometric limits of the LEM adapter.In this concept, the CSM subsystems support the laboratory module functions for the120-day resupply period. Consequently, Concept II has a laboratory which is dependenton the DSM subsystems. The third configuration (Concept III) is similar to Concept nwith respect to the addition of the separate laboratory module within the LEM adaptersection; however, the Concept III laboratory was designed to carry its own subsystems;therefore, it is considered to be independent of the CSM subsystems. The subsystemsin the CSM-independent laboratory module were designed for 1-year continuous operationwith resupply of expendables by the Apollo CSM as required. Included as a part of theoriginal Extended Mission Apollo study was the establishment of detailed developmentplans including costs, schedules and manufacturing and facilities plans. Thesedevelopment plans were based on an integrated but noninterference relationship withthe Apollo program and on maximum utilization of Apollo technology, facilities, etc.The laboratory module, for example, utilized a large percentage of Apollo/Saturninterstage tooling.

SID 64-1860-14

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SPARK PLUG-tht Electronlci Division of GENERAL MOTORS CORPORATION

Results of the initial Extended-Mission Apollo studies revealed several additionalfactors that warranted further investigation. The required 120-day resupply cycle dic-tated the use of advanced subsystem concepts in several areas. For example, theexisting fuel cell electrical power system in the Apollo was replaced with a solar cell-battery power system to eliminate the excessive weight and volume penalties attributableto the fuel cells and associated expendables. Additionally, a molecular sieve was em-ployed in the environmental control system to eliminate the large weight and volumeattendant with using the existing Apollo lithium hydroxide CO2 removal system for 120days. Consequently, the limiting mission duration which might result from using onlycurrent Apollo subsystem concepts was not known. Also, although the feasibility ofextending the life of the current Apollo subsystems had been established in the initialstudy, specific techniques for accomplishing this life extension had yet to be determined.

For these reasons, further studies were initiated as an addendum to the Extended-Mission Apollo contract. The purpose of these additional studies was to define thedesign characteristics and the maximum earth orbital mission duration of the ApolloCSM, assuming restriction to use of existing Apollo subsystem concepts. This con-figuration was identified as Concept 0. Included in this study was examination of eachsubsystem to determine the life extensions possible through the addition of spares andredundancies. The study concluded that the earth-orbital duration capability of Concept 0was approximately 90 days (based on Saturn IB pay load limits).

After review of these findings, NASA initiated the current Extended Apollo SystemsUtilization Study, one part of which is to define in depth the design and operationalcharacteristics of a vehicle based on the identical subsystems approach employed inConcept 0, but limited arbitrarily to a 45-day maximum mission duration. This vehicle,identified as Apollo X, appears to represent the most logical next-step in mannedspacecraft duration capability beyond the current Apollo and Gemini programs.

As part of this same study program, the characteristics of the CSM configurationsassociated with the dependent and independent laboratory module (as in Concepts II andIII) are being investigated relative to earth orbital and lunar missions of greatly extendeddurations. Under these Prolonged Mission studies, the Concept II CSM is identifiedas a Mission Support vehicle, since it supports the laboratory (AORL) by resupplyingsubsystems, crew members, and expendables. The Concept III CSM is identified asa Logistics Support vehicle since it resupplies only expendables and crewmen, with theCSM subsystems remaining in a quiescent state after docking to the laboratory module.The laboratory modules are being investigated separately (i. e., by another contractor)under a current MSC study program. The definition of the characteristics of theApollo X laboratory modules is also included as part of this separate AORL studyeffort.

The primary objective of the Apollo X study has been to define a standard space-craft design capable of alternately performing ex tended-duration lunar or earth-orbitalmissions of NASA near-term interest. Included in this primary objective were studies

-5- SID 64-1860-14

Thl» document contains tnfoithe meaning of the espionage lor revelttloD of which In any

.-.ing the national defense of the United States within; 18, U.S. C. Sections 793 and 794, the transmission

~^ unauthorised person Is prohibited by law.

SPARK PLUG-the Electronics Division of GENERAL MOTORS CORPORATION

to determine specific modifications required for each subsystem to accomplish theextended missions. Subsystem qualification test programs, which would substantiatethe analytically-derived extended-life subsystems, were also defined. In the experi-mental area, emphasis was placed on definition and integration of experimentalpackages (for earth-orbital missions) based on experimental lists provided by NASAat the initiation of the study. As an aid to NASA evaluation of the desirability of or require-ment for, the laboratory module, studies were also performed defining developmentfactors based on CSM changes attendant with and without the laboratory module. Alldevelopment factors investigations were based on an integrated, but noninterferencerelationship with the current Apollo program.

The Apollo X study was initially focused toward consideration of earth-orbital ex-perimental missions of up to 45 days' duration; spacecraft and subsystems requirementswere to be developed through the determination of the experimental mission require-ments. Secondary studies were aimed at determining the modifications required to the45-day earth-orbital CSM design when applied to a lunar survey mission (including upto 28 days in lunar polar orbit). Several weeks after study initiation, however, NASAdirected that equal emphasis be given to the lunar mapping mission, and requested thatprimary emphasis be placed on the definition of a standard CSM suitable for both thelunar and earth-t>rbital missions. Accordingly, a reevaluation was made of previouslydeveloped earth-orbital mission requirements. This reevaluation indicated that thelunar survey mission imposed overriding requirements in most subsystem areas.

As previously indicated, the overall study approach was based on derivation of astandard multimission CSM applied to different vehicle configuration arrangements, asindicated in Figure 2.

In the basic configuration, experimental work space is made available in thecommand module by elimination of the third crew member and couch, and by providingfor in-orbit storage of the center couch under the pilot (left-hand) seat position. Thisconfiguration is adaptable to earth-orbital experimental missions, but is not suitablefor the lunar missions since it would be required that the service module be completelyfilled (with propellant, fuel cells, cryogenic tankage, etc.) with no remaining volumeavailable in the SM for mission-oriented equipment. The basic configuration has, how-ever, the definite advantage of being able to provide an experimental laboratory forearly flights (such as biomedical and human factors experiments) without schedule de-pendency on a separate laboratory module. Additionally, these early flights wouldresult in a high-confidence base for performing subsequent missions with a laboratoryor lunar survey mission module.

The laboratory configuration has increased volume availability as indicated inFigure 2, through the addition of a separate laboratory module carried within the LEM 'adapter section. The separate laboratory module (approximately 1200-1500 cubic feetvolume) is sized by the provision of space inside the LEM adapter for a second moduleof identical geometry. This configuration can be applied to all of the earth orbital andlunar missions of current interest since mission-peculiar payloads can be installed

-6- SID 64-1860-14

Thii document contain* Info rmattojia&ict Ing the national defame of the United State* withinthe meaning of the espionage lawB,J^me 18, U. 8. C. Sections 793 and 794, the traoimlulonor revelation of which In any .dSr&nnersJo an unauthorized penon !• prohibited by law.

SPARK PLUG-the Electronic! Division of GENERAL MOTORS CORPORATION

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SPARK PlUG-tnt Electronic. Division of GENERAL MOTORS CORPORATION

within the laboratory module. The configuration is extremely versatile in that the ex-perimental volume availability can be controlled by the addition of either one or twolaboratory modules. In addition, a single module can be mounted on the LEM descentstage adapter, thereby permitting additional payload for lunar missions, orbital changecapability, or for providing a lunar landing cargo module.

All technical studies were predicated upon the basis of the Apollo Block II configura-tion. Since past studies revealed that subsystems characteristics materially influencedconfiguration design, heavy emphasis was placed on subsystems definition as Indicatedin Figure 3. In this regard, not only primary subsystem components were examinedin detailf but secondary components were also fully analyzed. In certain instances whereit was found that the addition of spare or redundant Apollo subsystem components mightnot suffice for the extended-life application, product improvement areas were identifiedand tentative solutions were established.

Previous studies had identified in-flight maintenance as a major study problem areain view of the number of spares and redundant components used for subsystem lifeextension. Correspondingly, in-flight maintenance was examined with regard to pos-sible methods of implementing these spares and redundancies as well as determining theoverall demands on crew time. These requirements were integrated into crew andexperimental station-keeping activities and served as a constraint in the experimentalscheduling area. In the experiment design and packaging analyses, it was possible toexamine the biomedical and behavioral measurements in great detail through the use ofbaseline data established in the initial Extended-Mission Apollo study. It is believedthat the Apollo X biomedical and human factors experiments as now defined are suf-ficiently detailed to permit the establishment of equipment specifications.

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-8- SID 64-1860-14

Thl> document contain* Inlormfet^nIffectlng the national dofenae of the United gtatea withinth. tanning of the aeplonago laweTrBJeie. U. 8. C. Bectlona 793 and 794, the tranamlaalonor revelation of which In any manqer^p^ an unauthorized peraon !• prohibited by law.

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SPARK PlUG-tht Electronic! Dlviiion of GENERAL MOTORS CORPORATION

I. INTRODUCTION

PURPOSE

This report details the results obtained by AC Spark Plug for use by S & ID toselect the^ optimum Guidance and Navigation (G & NJ|_System for use during the extendedApollo (Apollo X) lunar polar orbit mapping mission./ It also applies the defined G & NSystem to the earth polar and low inclination orbital Apollo X missions.

SCOPE

Constraints set by S & ID limited this study to the evaluation of the Apollo Block IIGuidance and Navigation System applied to the extended Apollo X missions and to thedefinition of means of system life extension without major system modifications. Thisreport describes the system in terms of equipment function, mission operation, relia-bility and maintenance, and physical parameters. Parametric studies have also beenincluded for the defined missions. This study excludes consideration of the ApolloGuidance Computer, which is covered in a separate report.

CONCLUSIONS

^74^This study shows that the Block n Apollo Guidance and Navigation System can provide

the lunar orbit navigation and local vertical stabilization requirements. Under the con-straint of minor modification of the existing Block n G & N, it is also shown that theApollo Block n Guidance and Navigation System can provide the apportioned reliabilityof . 9975 for the defined lunar polar orbit mission when the minor system modificationsare made and 91 Ib of spares and redundancy are added. '

The study also illustrates that the additional equipment required is very sensitiveto the particular level of reliability assigned to this study. The present optics reliabilitybarely exceeds that allocated to the G & N (less computer). Since the optics was notconsidered sparable (and not subject to major modification) the remainder of the G & Nassembly reliabilities were required to approach unity. In addition, approximately137. 5 hours of a total 157 hours of mission usage time accumulated by the G & N InertialMeasurement unit (in the Apollo X mission) are for providing local vertical stabilizationduring lunar polar orbit as opposed to a total usage time of 33 hours for basic guidancein the present lunar orbiting and rendezvous mission. These combinations of require-ments manifested themselves in two major areas; sparing of the IMU, and a redundantset of electronic coupling units. fr"*i,~_/"*L

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This document contain* info:the meaning of the espionage laws,or revelation of which In any

affecting the national defense of the United State* withine 18. U.S.C. Sections 783 and 794, the transmission

an unauthorised person la prohibited by law.

SPARK PlUG-tht Electronics Division of GENERAL MOTORS CORPORATION

The inherent reliability of the existing Block n G & N (less guidance computer) withminor modification and zero pounds of spares/redundancy in providing basic guidancefrom earth launch through establishment of lunar polar orbit and return to earth (withfull backup autonomous navigation) is . 990. Optimization of the various circuitry andreliability studies of the presently evolving Block n design are expected to improvethis basic guidance capability further. Future tradeoffs between related C/M subsystemsand the various backup and alternate capabilities they represent, such as the G & Nand SCS, in providing vehicle stabilization may permit alleviation of apportionmentsand usage times in both areas. £^1

n

-12- SID 64-1860-14

This document contain* Infothe meaning of the espionage lawaor revelation of which In any

affectlng the national defense of the United States withinitle 18> U. S. C. Section™ 793 and 794, the transnUsolon

an unauthorlted peraon Is prohibited by law.

SPARK PLUG-the Electronics Division of GENERAL MOTORS CORPORATION

II. REQUIREMENTS AND DEFINITION

REQUIREMENTS ANALYSIS

MISSION PROFILES

The three basic mission profiles considered for the Apollo X study were the LunarPolar Orbit Mapping Mission, the Earth Polar Orbit Mission, and the Low InclinationEarth Orbit Mission. Because the lunar polar orbit mapping mission is the longest andthe most complex, it was used to establish the spares/redundancy requirements and toidentify the critical components and subassemblies of the Apollo Block n G & N System.Then, the capability of this system design was used to determine its adequacy in theother two mission profiles.

The mission profiles are defined in Table 1, and the local vertical stabilizationcycle to be used during the lunar and earth polar orbit missions is shown in Table 2.These mission profiles are further detailed in the Appendix.

RELIABILITY REQUIREMENTS

The required reliability of the Apollo G & N System, exclusive of the guidancecomputer, is listed in Table 3.

DESIGN CONSTRAINTS

Certain constraints have been imposed upon the G & N System design so that anadequate system may be developed for the defined Apollo X missions with a minimumof time and cost.

Mission

Since the lunar polar orbit mapping mission is the most severe application of theG & N System in the Apollo X study, the system configuration must meet all thereliability and performance requirements of this mission. This same configuration willbe used for the two earth orbital missions.

The primary function of the G & N System in these missions is guidance. Naviga-tion is the prime responsibility of ground tracking,except for initial lunar orbit deter-mination where a combination of ground tracking and G & N System navigation can beused. None of the navigational capabilities of the G & N System will be deleted, how-ever, and thus it will serve as a backup to the ground tracking system.

~13~ SID 64-1860-14

This document contain* InfonnatfeQsgtfectlng the national defense of the United Statea withinthe meaning of the espionage laws, TltfteJS, U. S. C. Sections 793 and 794. the transmissionor revelation of which In any mannen, tdNan unauthorized person Is prohibited by law.

SPARK PLUG-the Electronics Division of GENERAL MOTORS CORPORATION

Table 1. Apollo X Mapping Mission Profiles

Mission Phase Time (i fc N Functions

Lunar Polar Orbit Mapping (Profile 1)

Launch and Parking OrbitInjection

Earth fjrblt

Trannlunar Injection andPlane Change

Trnnnlunar CoBHt

lAinar I'olar Orhlt

Trnnsoarth Injection

TranBuarth Const

Plane Change

Trannenrth Connt

Kntry

12 mlnulcB

:i hourH

fi mlnutefl

72 honr-H

2N (IH.VH

.'I mlnuton

1 hour

'i mlnutOH

71 houi'H

IH inlnuti'H

Monitor launch, compute nbort tra jeetorleH, Mini take over guidance.If necipSHsry.

Backup navigation.

Monitor IhriiHtlnK, compute abort trn|i<etorleH, and take over giilo1-anco If necessary.

Mldcourso AV corrcclloiiH and backup navigation.

OrhlUil norrncllonn ilurlnn flrNl f<iw o'rhlt.8 and locnl vorllcnlHtflhlllzatlon durlnR mapping. Corrncit luniir pulnr orhlt to80nm i 1. Inm (1"). Slnhill/.c CSM-l.»h to trunk localvitrllcnl within K). 12!i. (litRrnc. duaillKinil, all axiiH • 0.02 d<!K/Hiicdrift raid within doiiillmno.

(luldf) Bpacecrntt.

Backup navigation,

(iuldu Bpaoocrafl

Mldcourno AV uorrr.ctlonn and haidoip navigation.

HpacccnifL.

Knrth I'olar Orhlt (Profile

Launch and Parking Orhltln|«ctlon

Parking Orhlt ( lOOnm)

Tr.-m«f«»r Injection

Tr;inn|i:r TrBjoctory

K.irth Polar Orlilt ln]«cttlon

C.n-tli Polar Orlilt

I k - l l l l O H t

Knlry

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'i mlnutOH

4.'i diiyH max.

IH HcccmdH

14 to 21 mlnuton

Monitor launch, compute abort triilitrlorlitH, and tnkn ovor guldsuccIf nocossary.

Backup navigation.

Monitor thniKttiiK, coinpuU; itlwirt lru|«M'l<M li'H, iiml tnko over Kutd-anc.n II necuHHary.

Itacktip navigation.

Guide spacecraft. I'lrcularl/.e nrliit In :!l)l)nni i Hliini.

Orbital eorro.c.tlonn,loc.al vertical Hlulilll/.allon. Hiu'kup ui<vlKiitl«n.

CJuld(! Bpacocralt.

Ouldo Hpacecrall.

ll!w tncllnntlon Karth Orhll (I'rofllo :i)

(i2.'l NorrondHl.;iunch ami Parking OrbitInjection

Parking Orbit (80 nm)

'li;inHli;r Injection

Trannfor Trajoctory

K'iirth Orbit Injection

K'arth Orbit

Ik-hooxt

Kril ry

i)l) mlniiten

10 HficondH

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IH secondH

14 to 21 minutes

Monitor launch, compute abort trajectories, mid take over guidanceif necHHHiiry.

Backup

Ciuldu Hpucocraft.

Backup navigation.

(Juldo Rpacocraft. Circularly.!! orlilt In 20(1 nm i lOiun.

Orbit correction ond (', H, N checkout I « lorn entry. (Correct orbitafUjr 22 dayn to 200 nm i

Oulrte spacocraft.

spacocraft.

I—>: i

-14- SID 64-3860-14

Thll document conUlu Intormoftbu^fecllng the nutlonoj defenia at Uw United BUtM withinthe mMOlni of the e«plon«go Uw«T"W^to_ 18. U.» C. Seotlona 793 ud 794, the trinimllllonor revelation of which In «ny manner^o^ «n unauthorlied pereon In prohibited by law.

SPARK PLUG-th. Electronics Division of GENERAL MOTORS CORPORATION

rY

Table 2. Lunar and Earth Polar Orbit Mission Local VerticalStabilization Cycle

Operation

Start at Pole

Mapping Interval

No Mapping

Mapping Interval

No Mapping

Mapping Interval

No Mapping

Mapping Interval

No Mapping

Elapsed Time

Hours

00

1

5

0

5

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5

0

5

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00

00

15

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15

00

15

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15

Cumulative Time

Hours

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1

6

6

12

13

18

18

24

Minutes

00

00

15

45

00

00

15

45

00

Cycle to 82. 5 hours of mapping.

Mission

Lunar PolarOrbit Mapping

Earth PolarOrbit Mapping

Low InclinationEarth Orbit

Table 3. Mission Probability of Success

Probability ofSuccess

.99779

.99908

.99988

Remarks

This requirement was for the 28-day flight mis-sion. A later requirement has reduced this re-quirement to a probability of success of .9975.

It was desired that the maximum number of sched-uled mapping cycles be determined, so that thereliability requirements would be met afterentry. This requirement has been changed to thedetermination of reliability versus time forseveral different mapping cycles.

This requirement was for a 45-day mission. Alutor requirement was to determine what over-all reliability could be obtained for the 45-daymission using the G & N System designed to meetthe lunar polar orbit requirements.

-15- SID 64-1860-14

Thla document contain*the mrftnlig of the espionage law*,or revelation of which In any mannerll

ing the national defense of the United State* within, U.S.C, Sections 793 and 794. the tram mi* a I on

unauthorized person It prohibited by law.

SPARK PLUG-the Electronic* Division of GENERAL MOTORS CORPORATION

To account for the additional stresses experienced by the equipment, the equip-ment failure rates during mission thrusting phases will be multiplied by the following -|"K" factors. 1

Mission Phases "K" Factor '-|

Launch 3—-i

Translunar Injection and Plane Change 2 j

AV Corrections 2

Deboost • t 2 |

Transearth Injection and Plane Change 2

Entry 2 . j

Coast 1

Equipment Deenergized 0

G fc N Modification

The G & N System will conform to the Apollo Block II System as closely as pos-sible. Only minor system modifications are to be considered, and none will involvemajor G & N or spacecraft redesign.

Spares/Redundancy and Weight/Volumetric

System improvements that will not add weight or volume to the system will beexplored before considering additional equipment. Spares and/or redundancy deemednecessary to meet the required mission reliability and performance will be kept toa minimum in order that the additional weight and volume penalties be minimized.Sparing is considered preferable to redundancy because of mechanization and packagingconsiderations. The final selection of spares/redundancy should include considerationof a maintenance analysis of the equipment.

ASSUMPTIONS

Certain mission and equipment assumptions are required to proceed with the study.Some of the more important ones are:

• On the earth orbital missions, the first orbit may ho 'UHW! Tor orbit vorU'icntionand corrections.

1

-16- SID 64-1860-14

This document conUlnrettviniUon affecting the national detente of the United Slates withinth. m«.T>ing of the eaplonageTJjafcTltle IS, U.S. C. Sections 793 and TM, the tranimladonor revelation of which In any^SSMitetoan unauthorized person it prohibited by law.

SPARK PLUG-tin Electronic Division jUjEHUdl MOTORS CORPORATION

• Ground tracking and star occultation will be used for lunar orbit determination,and landmark tracking and star elevation measurements will only serve as backup.

• Minor computer program modifications will be made to enable the local verticalhold to be accomplished.

• The Apollo Block n G & N System configuration (presently still under design anddevelopment) will not change appreciably from the current configuration.

• The stressed module reliabilities, as defined by MIT/IL for Block I, areapplicable to similar Block II modules. Reliabilities are based on averagefailure rates and parts counts where stressed failure rates do not exist.

• Launch time is time zero for all part and module reliabilities.

• Sparing cannot take place 1 hour before deboost to lunar orbit, 1 hour beforeentry, or during any thrusting phases. The times prior to all other thrustingphases were not considered critical since thrusting could probably be postponedif sparing were required.

ENVIRONMENTAL CONDITIONS

The basic environmental conditions for the lunar Apollo mission have been sum-marized in NAA Specification MC 999-0051, Apollo Environmental Design and TestRequirements, and in NASA Document ND 1002037, Apollo Airborne Guidance andNavigation Equipment Environmental Qualification Specification.

The Apollo X G & N equipment is expected to be subjected to the same conditionsduring a flight as shown in the above specifications,with the following exceptions.

• Time spans of exposure to various dynamic and climatic environments will belengthened in some instances.

• The spacecraft crew compartment and various portions of the G & N equipmentwill be subjected to an atmosphere composed of 50 percent nitrogen and50 percent oxygen at 7 psia and 70° F over the entire mission.

G & N equipment design is not expected to be affected by the above environmentalchanges.

_17_ SID 64-1860-14

This document contain* InfoPhi^tan affecting the national defenee of the United States withinthe meaning of the eiplonage lawi^SfelglB. U.S. C. Section. 793 and 794. the tranemlulcnor revelation of which In any mannSrlb^jn unauthorized penon t> prohibited by law.

SPARKPUJ&^iheElertren|« Djvlslon^ o| GENERAL MOTORS CORPORATION ._

BLOCK H G & N SYSTEM DEFINITIONr—\

The Apollo G & N System is capable of independently performing two basic functions: 'guidance and navigation. To accomplish the guidance function, three accelerometersare orthogonally mounted on a gyroscope-stabilized platform that is isolated from Ispacecraft angular motion by a gimbal system. These accelerometers sense velocity l

changes, while the gyro-stabilized platform and gimbal system senses changes in the ^spacecraft attitude. Together, they provide steering and thrust control information. jA sextant and a scanning telescope are provided to take sightings on celestial bodiesand landmarks. These sightings are used to determine spacecraft position and velocity _^and to establish a precise alignment of the stable platform. |

A digital computer serves as the primary data processing element of this system. _It contains a catalog of celestial objects, and is programmed to calculate steering and • \thrust commands using information obtained from optical sightings or ground trackingstations. The G & N System together with the Reaction Control System form a loop .-,that controls the spacecraft attitude and translation. A second loop, formed by the :' jG & N System and the Service Propulsion System, provides control of spacecraft thrust-ing. -,

iThe Apollo Block n G & N System is divided into three major subsystems: inertial,

optical, and computer. A functional diagram of this system is shown in Figure 4. The /•>three subsystems, or combinations of subsystems, can perform the following functions: \

• Calculate the position and velocity of the spacecraft by inertial measurementsand by the following optical navigation sightings: j

Landmark trackingStar occultations ~iStar — horizon sightings • jStar—landmark sightings

• Store and operate on data received from ground tracking, I

• Align the stable member to an inertial reference,using precise optical sightings,

• Calculate necessary velocity corrections and provide control of spacecraft thrust- «.ing to effect the required trajectory, '

• Control the spacecraft attitude with respect to an inertial reference, •"*;

• Provide backup to the Saturn IV B (SIVB) booster guidance system,

• Provide the astronaut with a data display indicating the status of the guidance "~]and navigation problem. j

~18~ • SID 64-1860-14

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Ttui document conuliu lutormirtiw&ectlng the national defenie o/ the United State* withinthe meanlnf a) the elplonage lawj/Tltte 18, U.S. C. Sections 793 and 794. the tranimliilonor revelation of which In any mannrS an unauthorized perion It prohibited by law.

SPARK PLUG-the Electronics Division of GENERAL MOTORS CORPORATION

CO N P rB'^MJHA^ "1INERTIAL SUBSYSTEM

The Inertial Subsystem (ISS) is used to establish and maintain an inertial refer- Ience that provides the capability to measure linear acceleration and spacecraft attitudewith respect to inertial space. '""]

Inertial Subsystem Equipment Definition ' .-^- . ' • - . ' • ' |

This subsystem consists of an inertial measurement unit (IMU), three electroniccoupling units (ECU), and the major portion of the power and servo assembly (PSA).. _,The IMU contains a three-gimbal inertial platform containing three stabilization {gyros (25 IRIG Mod 2), three accelerometers (16 PIP A), and the associated electronics.The IMU also contains a simple on-off temperature controller, which is used to main- ,-^

I tain the 16 PIPA's and the 25 IRIG's at their proper operating temperatures. A thermo- Istat is used to monitor the temperature of the stable member. When a high- or low-temperature condition exists, the thermostat will issue an alarm signal, which is dis- ---5played by means of a light. \ \

The ISS portion of the PSA contains the electronics and power supplies necessary —{

to support the IMU. Figure 5 indicates the major elements comprising the inertial 1attitude reference portion of the ISS.

'•—1Linear accelerations along three orthogonal reference axes are sensed by the three |

accelerometers. Using time reference pulses from the computer, the sensed accelera-tion is integrated, and information in the form of velocity changes is sent to the com- nputer for use in determining spacecraft velocity and position. A block diagram of a !16 PIP A loop is shown in Figure 6.

•sThe ECU's have two basic functions: (1) converting analog signals representing . j

the IMU gimbal angles to digital signals for the guidance computer, and (2) convertingdigital signals from the computer to analog commands to coarse align the IMU gimbals. ~\In addition, the ECU's can also provide analog attitude error signals to the SIVB as a \backup to the SIVB guidance and attitude error information to the astronaut. A blockdiagram of one ECU is shown in Figure 7. """•

\

L'lertial Subsystem Modes of Operation -^. _ . . j• i

The Inertial Subsystem has two basic states, standby and operate. In the standbystate, only temperature control and suspension for the inertial components are main- ~~,tained. In the operate state, the inertial subsystem is in one of the following four modes, , ]which are selected by the guidance computer.

-20- s^___ SID 64-1860-14

nit document contain* lnformatlofiWlec«j>£»he national defeneo of the United Statn withinth. meaning at the eeplonage lawa^W^SBTu. S. C. Section! 793 and 794. the tranimlulonor revelation of which In any'^MnneY to an unauthorized perion It prohibited by law.

SPARK PlUG-th. Electronic. Dlviiion of GENERAL MOTORS CORPORATION

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Thli document conulnt InformiliVlltectlng the national defeni* at the United SUtee withinthe meaning ol the eeplontgo lawi, TMle 18, U.6.C. Section. 793 and 794. the trmemleolonor revelatloo ol which In any maraj^jo an unauthorlttd peraon la prohibited by law.

SPARK PLUG-th* Electronic* Division of GENERAL MOTORS CORPORATION

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rating tha nnllun»l dcfon** of the Uttltad §UU« withinin. u.H. C. Hootlona 7W and 704. lh« trwi>ml»lan

unauthorlied parian !• prc^lbltMl by law.

SPARK PlUG-thi Electronic* Division of GENERAL MOTORS CORPORATION

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SID 64-1860-14

This document contAloj in^o'c^aHpn affecting the national defense of the United States withinthe meaning of the e§pionag?TOT«jtM.le 18, U. S. C. Sections 783 and 794, the transmissionor revelation of which In any mSgaSSgWi"! unauthorized person is prohibited by law.

SPARK PlUG-thi Elettroniu Division of GENERAL MOTORS CORPORATION

Zero ECU—-,

In the Zero ECU mode, the gimbal angle counters in each of the three ECU'S and ^in the computer are reset to zero and A0g pulses are inhibited from being transmittedto the computer. The stable member attitude is maintained by the 25 IRIG error sig- ~jnals. i

Coarse Align j

In the Coarse Align mode, the guidance computer commands the IMU gimbal to a ._,desired orientation. The computer calculates the differences between the present and . ithe desired gimbal angles. These differences are transmitted to the ECU'S, where

» they are converted to analog error signals. The gimbal servos use these error signals _to drive the gimbals to the desired orientation. When a gimbal angle changes by an ; Iincrement of 40 arc-seconds, a pulse is transmitted to the computer by the ECU.A functional diagram of the Coarse Align mode is shown in Figure 8. r*»

Fine Align

In the Fine Align mode, the computer commands a "fine" change in the inertial 'orientation of the stable platform. The guidance computer commands the pulse torqueelectronics to torque the floats of the 25 IRIG's from null through a small correction ~~langle. This produces an error signal to the gimbal servos, which change the inertial 1

orientation of the stable platform such that the 25 IRIG floats are again at a null. FineAlign is the only mode in which the IRIG floats can be torqued since in the other three jISS modes, the pulse torque electronics are deactivated. A functional diagram of the 'Fine Align mode is shown in Figure 9.

Attitude Control

The Attitude Control mode is identical, as far as the ISS is concerned, to the FineAlign mode with the exception that the pulse torque electronics are deactivated. Sincethe platform is inertially stabilized and the ECU'S transmit changes in gimbal anglesto the guidance computer, the computer can calculate the attitude of the spacecraft withrespect to the inertial reference. The PIPA loops are in operation whenever the ISSis in the operate state. However, under normal inflight conditions, the outputs fromthe PIPA loops to the computer will only be meaningful when the ISS is in the AttitudeControl mode. . .

• i

-24- SID 64-1860-14

Thli document contain* Inforftwjtoiflalfectlng ^e national defenae of the United Btatea withinthe meaning of the espionage lawa^1M4£l8, U. S. C. Section! 793 and 794. the tranamlialonor rorelatlon of which In any matner^to. an unauthorized pereon la prohibited by law.

SPARK PLUG-the Electronics Division of GENERAL MOTORS CORPORATION

ISS AGC

PresentGlmbalAngle

Error inGimbalAngle

Figure 8. ISS Coarse Align Mode

ISS AGC

p

nn

SpacecraftNon Field — — fc-

Accelerations

SpacecraftAttitude *

AttitudeError Sigs.Tn STVB ^Analog

and/orAstronaut

Display

PIPALoops

InertlalAttitude

Reference

Pulses

_ Pulses

(Analog)GlmbalAngleSignals

i r

ElectronicCoupling

Units

j

Pulses

Pulses

Non- Field

Velocity Changes

Inertlal Referenc

(Fine Align Mode

Spacec raft

Changes

Attitude ErrorSignals

eaOnly)

IFigure 9. ISS Attitude Control and Fine Align Modes

r-25- SID 64-1860-14

Tbti document aontmlu (nforthe meaning of the eapkmago UWBor revelation of which In any mann

natlnnoJ <Jcfon«e of the t/alted BUtea within18, U.B.C. K ceil on* 793 and 794, the truicmUftlon

unauthorized per ion u prohibited by Uw.

SPARK PLUG -Id* Electronic Division of GENERAL MOTORS CORPORATION

OPTICAL SUBSYSTEM•— i

The purpose of the Optical Subsystem (OSS) is to provide instruments with which ithe astronaut can measure angles between: (i) a star and the horizon, (2) a star anda landmark, (3) a star and the navigation base, and (4) a landmark and the navigation -jbase. Since the IMU is precision mounted to the navigation base, the guidance com- 1puter can combine the information received from both the ISS and OSS to perform opera-tions necessary to the mission success. ~]

i

Optical Subsystem Equipment Definition _^

The OSS consists of a scanning telescope, a sextant with a star tracker and ahorizon photometer, and their associated electronics and controls. The sextant is a ^dual line-of-sight, 28-power, 2-degree field-of-view instrument. The star line-of- jsight (StLOS) can be articulated 55 degrees from the landmark line-of-sight (LLOS)about the trunnion axis, which in turn, can rotate ±270 degrees about a shaft axis, .->.which is coincident with the LLOS. The star tracker is used to keep the StLOS pointed jat a preselected star while the horizon photometer, which is parallel to the LLOS, ispointed at the "blue line" of the earth by controlling the attitude of the spacecraft. The n

scanning telescope is a single line-of-sight, 1-power, 60-degree field-of-view instru- I jment. Its LOS is fully articulated about the same polar coordinates as the sextantStLOS. The available view is limited to approximately a 55-degree, half-angle cone — jdue to spacecraft structure. . j

The StLOS can be controlled by the astronaut's positioning of the hand controller nstick. The hand controller supplies a voltage to an integrator that drives the sextant jand the scanning telescope, which is slaved to the sextant. Angular position of theStLOS is transmitted to the guidance computer by means of electronic coupling units ~)(identical to the inertial subsystem electronic coupling units), which encode the angular .1rotations of the Ix and 64x resolvers on the trunnion axis, and the Ix and 16x resolverson the shaft axis. The StLOS can also be controlled by the computer via the electronic ncoupling units. , }

Optical Subsystem Modes of Operation ; j

The Optical Subsystem has three prime modes of operation: Zero Optics, Manual, _,and Computer. These modes and their submodes are discussed below. !

-26- - SID 64-1860-14

Thli document oonUIn* Intormifhaygjoctlrig tho nutlonil d<i(on» at Ui. United SUtu withinDM ntMJilnf of Uu> »pla»«o !«»«, rftffcj^ U.B.C. S.ctlon. 7M iod 7»4. Die truunUMIcnor revoltllon of which In uy minnan l8»»i uniuthorltod perion li prohibited by Uw.

SPARK PlUC-tht Electronics Division of GENERAL MOTORS CORPORATION

Zero Optics Mode

In the Zero Optics mode error signals are generated in the 64x and lx sextanttrunnion resolvers and in the 16x and l/2x sextant shaft resolvers. These signalsare connected to their respective motor drive amplifiers via two-speed switches. Inaddition, there is scanning telescope trunnion resolver follow-up between the two lxtrunnion resolvers, and as a result, all four servo channels are driven to zero. Whenswitched to the Zero Optics mode all optics ECU counters are cleared. See Figure 10.

Manual Mode

In the Manual Direct mode the sextant trunnion and shaft angles are driven directlyby the hand controller. Generally, the scanning telescope follows the sextant position.The Manual mode contains three scanning telescope options (0-Degree Offset, 25-DegreeOffset, and Star Line), two sextant options (Direct and Resolved), and an Automatic StarTracking Option.

0-Degree Offset. In this mode, the 800 cps reference is placed upon the cosine windingof the scanning telescope trunnion lx resolver. This constrains the scanning telescopeto shaft motion only.

25-Degree Offset. The 25-Degree Offset mode is the same as the 0-Degree Offset,except that . the 800 cps reference is run through a 25-degree offset transformer tothe sine winding of the scanning telescope trunnion lx resolver. This offsets thescanning telescope LOS 25 degrees with respect to its zero position, thereby providingthe capability for visually scanning a 110-degree cone by rotating the shaft. This isshown in Figure 11.

.Star Line. In the Star Line mode, the scanning telescope trunnion is slaved directlyto the sextant trunnion by means of both resolver and tachometer follow-up.

Direct. Up-down and left-right motion of the hand controller results in polar coordinateimage motion. Moving the hand controller to the UP position results in image motionradially inward (that is, trunnion angle increasing). Moving the hand controller to theright results in counterclockwise image motion (that is, shaft angle increasing). TheDirect mode is to be used when incrementing the counters to a specified value deter-mined from the Map and Data Viewer.

_27_ SID 64-1860-14

Thle document oontalai InforirntYSnsgffSctlng the national defence of the United SUUa withinthe meanlaf of the espionage law*. T""m^iU. 8. C- Section! 793 and 794. the tranamlaelonor revelation of which In any manneMUcT!fe»4iinauthorlied perion It prohibited by law.

SPARK PlUG-tht Eltctronics Division of GENERAL MOTORS CORPORATION

Trunnion Axis

> LOS

Shaft Axis

1

1

1

1

Figure 10. Optical Subsystem — Zero Optics Mode

-28-

Thl« document eonUiiu lafarmaHon aftfctlng the national dafenia of the Unlt«d State* wtthlnth» mMalaf o< the atplocage^^^me 18. U.S.C. Section* 793 and TM, th« tranamlHlonor revelation at which ln^£B£^manne'r^4P &a unauthorized person !• prohibited by law.

SID 64-1860-14

Page Intentionally Left Blank

SPARK PUIG-thj Electron^ Division, of^ GENERAL MOTORS CORPORATION

IResolved. Up-down and left-right motion of the hand controller results in up-down andleft-right image motion. The resolved mode is to be used when attempting to super- ~|impose a landmark and star. See Figure 12. <

1Automatic Star Tracking. In this mode, a star tracker keeps the sextant StLOS aligned jto a star. This mode is unique in that it is the only mode in which the hand controllerhas the capability of directly driving the scanning telescope trunnion motor drive amplifier. ~-iThe shaft error goes to the sextant shaft motor drive amplifier via the cosecant amplifier. |See Figure 13. The Manual Track mode is to be used when taking star-horizonmeasurements. —j

t

Computer Mode

In the Computer mode the scanning telescope trunnion servo channel follows the . 'sextant trunnion by means of Ix resolver and tachometer follow-up. The Computermode contains three modes, which are described below. i

Computer Zero Optics "ji

The Computer Zero Cptics mode is the same as the normal Zero Optics mode withthe one exception noted above, that is, tachometer follow-up in the trunnion servo Hchannel. i

Computer Operate jj

In the Computer Operate mode the digital-to-analog converter output commands n

the sextant in both trunnion and shaft. ' i..' *

Tracker Operate ~~i

The star tracker output commands the sextant trunnion and shaft; the trunnion iscontrolled directly, and the shaft is controlled via the cosecant amplifier. ~]

I

COMPUTER SUBSYSTEM -,i!

The Computer Subsystem, discussed in detail in Volume 16 of this study, consistsof the Apollo Guidance Computer (AGC) and two display and keyboard (DSKY) units -jmounted on the D & C panel. The two DSKY's enable the astronauts to communicate jwith and control the computer. The DSKY's can enter information into the computer'serasable memory, display information present in the computer, or initiate programs. ~]

-30- SID 64-1860-14

Thla domimgnt contain* Information affatting the national defense of the United Btatei withinthe meaning of the etplonage-TS^KT^ft 18, U.S. C. Section* 783 and 794. the truumJMfonor revelation of whloh In, any^jnann>»^Q an unauthorized pereon !• prohibited by law.

SPARK PlUG-th. Electronics Division of GENERAL MOTORS CORPORATION

HandController

Stator

Rntn

CosecantAmplifier

64x lx

rSCTMDA

MotorGenerator

1

5, 952^

^S 1.> LOS

IJJ

Compensation

LOS

>LOS

Figure 12. Optical Subsystem — Manual Resolved Mode with Scanning TelescopeFollowing St LOS

-31- SID 64-1860-14

Tbi* document eonUiM tA/orm£Rl^gf0etlng th0 n*tlon«J d«/en*« at ttt» Uafttfd St«to0 wlt&lath« munlnf of the ••plondge liw«/Tl"^|j<J- S. C. Soct(on» 783 and TM, the truamlxlonor r«vol«tl<jn of wWoh In «ay m«raer ! Tto uiiutliorlixl p«non i< prohlbltad by l«w.

SPARK PUIG-tht Electronics Division of GENERAL MOTORS CORPORATION

Star TrackerElectronics

1

1

1

1

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^-

r*SCTMOA

MotorGenerator

s.asa/^jS 1

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rr/1/ tri

* *> ~r

Figure 13. Optical Subsystem — Tracker Operate —Manual Optical Subsystem

-32- SID 64-1860-14

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The AGC is the central processor for the G & N System. It is also the clock orbasic time and frequency reference of the spacecraft. The AGC can communicate withthe sextant and scanning telescope via"the ECU'S and can also communicate with theastronaut via the DSKY's. In addition, the AGC can count pulses from the PIP A loopsand read the IMU gimbal angles. The AGC can send information to earth viatelemetry, and receive telemetry information on an uplink. During guidance modes,the AGC can control and stabilize the spacecraft and start and stop the service moduleengine.

n

n

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MISSION OPERATION _-

Operation of the standard Apollo Block II G & N System may be best illustrated bydescribing its use in the Lunar Mapping Apollo X mission. This description mainly ~]treats the functional capabilities of the G & N System that are planned to be used. The Imajor phases in the mission are: n

• Launch and Parking Orbit Injection, ^

• Parking Orbit, _^

• Translunar Inject, .' j

• Transposition and Docking, 1 t• Translunar, I

• Deboost to Lunar Orbit,

• Lunar Orbital Correction, < I

• Lunar Mapping, _^

• Transearth Inject, 1

• Transearth, ^

• Entry. j

LAUNCH AND PARKING ORBIT INJECTION S!

Prior to launch, the IMU is aligned to an earth coordinate system by the guidance ._^computer. The guidance computer fine aligns the IMU stable member such that two of ithe PIP A loops sense no acceleration due to gravity, and the 25 IRIG's sense a predeter-mined earth rate. At lift-off the guidance computer automatically switches to a boost —monitor program when a discrete is received from the booster (SIVB) guidance system. JDuring boost monitor, the ISS is in the Attitude Control mode and transmits attitudeand velocity information to the guidance computer. The computer then calculates the -,trajectory and attitude error of the spacecraft. The total attitude and the .attitude errors [are displayed to the astronaut. In case the actual trajectory deviates too far from thatdesired, either an abort sequence or a takeover of the SIVB guidance can be initiatedby the astronaut, by automatic onboard malfunction detection, or by a ground station. :When the parking orbit has been achieved, the SIVB engines are shut down.

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PARKING ORBIT

During the earth parking orbit phase, earth based ground tracking is used toaccurately determine the spacecraft position and velocity as a function of time. Thisinformation is transmitted from the ground to both the G & N and SIVB Guidance Systems.During this phase, the Inertial Subsystem remains in the operate state. Also duringthe parking orbit phase, the IMU alignment will be updated in the following manner. TheISS will be placed in the Zero ECU mode by the computer, so that the gimbal angle countersin both the computer and ECU's will be reset to zero. The computer then places theISS in the Attitude Control mode. Using the gimbal angles, the computer calculates theattitude of the spacecraft with respect to the degraded inertial reference defined by theIMU. The OSS is now turned on in the Computer mode. The computer places the OSSin the Zero Optics mode, and then commands the sextant star line-of-sight (StLOS) toan attitude at which a predetermined star should be located. The computer controlsthe attitude of the spacecraft during this time to allow accurate measurements. Theastronaut assures that the proper star is in the sextant's field-of-view by switching tothe Manual Resolved SCT Follow StLOS mode, and if necessary, slews the StLOS suchthat the star is centered in the scanning telescope field-of-view. The astronaut thenenergizes the star tracker, which "locks on" to the star. The computer records theIMU gimbal angles, the optics shaft angles, and the trunnion angles. In a similarmanner, a second star-sighting is performed. The computer calculates the alignmenterror of the stable member about its three axes, switches the ISS to the Fine Alignmode, and pulse torques the IRIG's to correct for the alignment error.

TRANS LUNAR INJECT

The inertial reference will be updated prior to translunar inject. The G & N Sys-tem will perform the same monitor function as it did for the launch to parking orbitphase of the mission.

TRANSPOSITION AND DOCKING

After translunar inject, the Command and Service Modules are disconnected fromthe SIVB structure, rotated through 180 degrees, and then docked to the LaboratoryModule. The Laboratory Module is then disconnected from the SIVB structure, whichis no longer needed. During this phase, the G & N System "holds" the attitude of thespacecraft when the astronaut is not commanding the spacecraft with the Rotational —Translational hand controller. When the hand controller is in its center position, thecomputer controls the Reaction Control System jets to keep the IMU gimbal angles fromchanging. Since the stable member is fixed with respect to inertial space, any changein gimbal angle will correspond to a change in spacecraft attitude.

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' \

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THANSLUNAR

During the major portion of the translunar phase, the navigation function will be 'performed by ground tracking, and position and velocity data as a function of time willbe transmitted to the computer, so that, in the event of failure of the communications !~|link, the spacecraft and crew can continue the mission or return to earth by using the 'navigation capability of the G & N System.

Three velocity corrections are nominally planned during the translunar phase.Prior to these corrections, an alignment of the stable platform will be performed inthe following manner: The ISS will be switched to the operate state early in the pro-cedure to allow adequate time for thermodynamic transients on the inertial componentsto die out. The ISS is switched to the ECU Zero mode and then to the Attitude Control *~^mode. ' I

' i

The OSS is turned on and placed in the Zero Optics mode. The astronaut examines ^the star field in the scanning telescope for a star that is contained in the computer star •catalog. If a catalog star is not found in the 60-degree field-of-view, the astronautswitches to the Manual Direct 25-Degree Offset mode and slews the sextant and scanning ,->telescope about the shaft axis. This enables the astronaut to search a 110-degree ifield-of-view without moving the spacecraft. Once a catalog star is found, the scanningtelescope and sextant are slewed until the star is located on the telescope center line. -,The Manual Direct StLOS mode is selected, causing the scanning telescope to repeat jthe sextant St LOS.

The hand controller is operated until the star is located at the center of the tele- jscope field-of-view. In addition, the star will be in the sextant field-of-view. Thestar tracker is turned on and the sextant "locks on" to the star. The shaft, trunnion, -vand gimbal angles are recorded in the computer. j

A second star is sighted in a similar manner, and the shaft, trunnion, and gimbalangles are again recorded. The computer calculates the stable member alignment error ;and the gimbal angle errors. The computer switches the ISS to the Coarse Align modeand commands the gimbals, via the ECU's, to the proper angles. The ISS is then -^switched back to the Attitude Control mode. i

Two more star sightings are performed in a similar manner. The computer cal- —;culates the error in the stable member alignment with respect to inertial space. The \ISS is switched to the Fine Align mode and the IRIG's are torqued to correct for thealignment error. .—)

. . .1

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" SPARK PlUG-thi Electronics Division of GENERAL MOTORS CORPORATION

Once an accurate attitude reference is established by the ISS, the guidance com-puter commands the spacecraft, via the Reaction Control System, to the proper attitudefor thrusting with the Service Module engines. Since thrusting will only last a fewseconds, no thrust vector control is programmed by the computer.

DEBOOST TO LUNAR ORBIT

Deboost to lunar orbit is accomplished in a manner similar to mid-course velocitycorrections. Since deboost to lunar orbit will last a few minutes, the computer, throughthe digital-to-analog portion of the optics ECU'S, will control the gimbals of the ServiceModule thrust engine to provide steering as well as thrusting control.

LUNAR ORBIT CORRECTIONS

Verification of the lunar orbit will be performed by ground tracking from the earthand/or by onboard star occultation measurements. Orbital corrections will be per-formed in the same manner as mid-course velocity corrections.

LUNAR MAPPING

The inertial reference will be accurately aligned to inertial space, and local verticalwill be determined and maintained by the G & N System with the aid of the Reaction Con-trol System. The computer will: (1) calculate the desired attitude of the spacecraftwith respect to inertial space as a function of orbital position, (2) compare the desiredattitude with the attitude defined by the IMU inertial reference, and (3) correct thespacecraft attitude to the desired attitude.

TRANSEARTH INJECTION

Inertial guidance of the spacecraft is performed by the G & N System in the samemanner as deboost to lunar orbit except that 1 hour after injection, a trajectory planechange is performed.

TRANSEARTH

The transearth phase is conducted in the same manner as the translunar phase.

ENTRY

Shortly before entry, the stable platform is aligned using optical sightings. Duringentry, the G & N System, by means of inertial guidance, controls the direction of thelift and drag forces by rolling the spacecraft. The G & N System is capable of landingthe spacecraft within a predetermined area on the earth to facilitate recovery of thecrew and Command Module.

_37_ SID 64-1860-14

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•2205^i**1*4 *

SPARK PUIG-ih. Electronic! Division of GENERAL MOTORS CORPORATION

III. TECHNICAL ANALYSES

BLOCK H PERFORMANCE ANALYSIS

Performance of the Block II G & N System was investigated in three main areas:the first was a consideration of the navigational accuracy; the second, the ability ofthe system to maintain local vertical; and finally, system reliability degradation withtime.

The navigational and local vertical stabilization requirements are considered onlyduring the lunar polar orbit phase. The remaining mission phases are considered tobe sufficiently similar to those existing in the lunar orbital rendezvous mission, andtherefore the G & N System guidance monitor, control, and autonomous navigationcapabilities are within the performance requirements specified?1 The capability of theG & N guidance and monitor functions necessary for injection into 200 nm circular earthorbits from initial parking orbits (Profiles 2 and 3) has been previously evaluated in theXMAS and MODAP studies?* For the loose tolerance required on orbit circulariza-tion, the G & N is considered more than adequate.

The G & N System ability to provide local vertical stabilization signals for thespecified periods during lunar mapping operations involves the establishment of aninertial reference with a known angular relationship to the initial mapping point localvertical. The computer (with minor program revisions) can then compute the requiredvehicle attitude with respect to the inertial reference as a function of time (based on aknowledge of orbital position and velocity), and can issue the appropriate attitude com-mands to maintain the vehicle within the desired attitude deadband and rate tolerances.

The basic Apollo Block n G & N System reliability analysis considers the applica-tion of the equipment as used in all the mission phases of the lunar polar orbit missionprofiles. The applicable Apollo Block II G & N System error sources are listed inTable 4.

* MIT/IL Performance Specifications, Guidance and Navigation System — Block II,10 June 1964 (Confidential).

**AC Spark Plug: Modified Apollo Logistics Spacecraft (MODAP) and Extended MissionApollo Space Station (XMAS) Missions Final Study Report, 15 December 1963.

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Table 4. Apollo Block n G & N System Error Sources

Sensor Error Source

Gyro

Bias DriftSA UnbalanceIA UnbalanceCompliance

Accelerometer

BiasScale FactorNonlinearitySensitive Axis Misalignment

Initial Alignment or Realignment

Prelaunch

In Space

Optics

Star-to-Horizon "blue" line(Sextant Tracker-Horizon Photo-meter combination for orbitnavigation)

Star-Landmark(Sextant —for translunar navi-gation)

Landmark Track

Star Occultation (An error sourcenot related to the system)*

la Value

1 meru10 meru/g10 meru/g1 meru/g2

1. 5 x 10"4 g6x 10~5g/g2x 10-5g/g2

20 arc-seconds

40 arc-seconds level6 arc-minutes azimuth60 arc-seconds all axes

6 arc-minutes

10 arc-seconds

6 arc-minute s

2 to 6 arc-minutes

. V

*Baker, D. S., et al: Lunar Orbit Determinsition by Star Occultations and MSFNTracking, MIT Report E-1429, September 1963.

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LUNAR ORBIT NAVIGATION

For lunar orbit navigation, it is required that altitude information be obtained witha standard deviation of less than the orbit tolerance of ±1. 1 nm. Ground tracking oronboard navigation (star occultation or landmark tracking) are considered as possiblenavigation modes. Parameters considered are measurement uncertainty, samplingtime, and initial orbit injection errors.

Using the scanning telescope for star occultation or landmark tracking, measure-ment uncertainty standard deviations of 2 to 6 arc-minutes were assumed. For groundtracking, range accuracy standard deviations of 50, 500, and 5,000 feet were considered.Reports indicate that range accuracies of 50 feet are attainable. * It is assumed thatno correlation exists between the errors in consecutive measurements.

The time between consecutive measurements strongly influences the time necessaryto obtain the necessary accuracy. Conservative estimates concerning the sampling rateof the ground tracking stations were made. This was further enhanced by the assumptionof a worst case as far as visibility of the space ship from ground is concerned, for ex-ample, it was assumed that the space ship is only visible for a half revolution. Ananalog philosophy was applied to onboard navigation.

Position and velocity errors at the time of injection into lunar orbit were chosenaccording to the performance specifications. However, to obtain a more conservativeestimate of the performance, errors twice as large were also considered. In addition,it was assumed that the data is processed in an optimum linear filter?

The results, as shown in a series of plots (Figures 14 through 22) of position andaltitude accuracy versus the number of orbital revolutions, indicate that the requiredaltitude accuracy of 1. Inm can be obtained after one revolution. From these curves,the following additional conclusions were drawn:

• Ground tracking is superior to onboard navigation,

• Changes in the injection errors have minimal effect on the measurement uncer-tainty after one revolution,

• If the onboard instrument accuracy for star oecultations is sufficiently good, theinterval between observations can be increased after one revolution,

• The main position error orientation is more tangential than normal with respectto the orbit.

*Baker, D. S. (op. cit.)Smith, L G. and E. V. Harper: Midcourse Guidance Using Radar Tracking and On-

Board Observation Data, NASA TN D-2238.

**Sorenson, H. W. , F. Salomon, T. Kido: Application of Wiener-Kalman FilterTheory to Orbit Determination, AC Spark Plug ACRD Report, August 1963.

-41- SID 64-1860-14

Toll document contain* Infth. mualof at the oiplonago laor rmlatlon of which In any

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Figure 19. Maximum Position Uncertainty (la) VersusGround Tracking Accuracies

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LOCAL VERTICAL STABILIZATION

The local vertical attitude deadband requirement for the lunar mapping mission isgiven as ±450 arc-seconds about all axes. Contributing factors to the error in localvertical attitude include the initial space alignment of the inertial platform, the gyrobias drifts, the spacecraft positional uncertainty in orbit, and errors in the vehicleattitude control system. The alignment uncertainty between the platform axes andthe star lines-of-sight has a standard deviation of 60 arc-seconds about any axis. Thegyro bias drift has a standard deviation of 1 meru (0.9 arc-sec/min). The spacecraftpositional uncertainty at the start of mapping is assumed to have a standard deviationof 0. 5nm. This positional uncertainty yields an angular error of 110 arc-secondsin an 80 nm lunar orbit. Vehicle control system errors are not considered here.

Since the above error sources were independent and random in nature, they werecombined by the root-sum-square technique to determine an overall standard deviation.See Table 5. Tabulating these error sources versus time, it is seen that the attitudedeadband requirements are met with a 0. 5nm position uncertainty. The additionalvehicle control system uncertainty allowable lies in the range of 50 to 60 arc-seconds.

Table 5. Local Vertical Stabilization Error (lo)

Time(min)

0

15

30

45

60

75

AlignmentUncertainty(arc-sec)

60

60

60

60

60

60

Gyro BiasDrift

(arc-sec)

0

13.5

27.0

40.5

54.0

67.5

S/C PositionUncertainty(arc-sec)

110

110

110

110

110

110

Local VerticalUncertainty(arc -sec)

125

126

128

132

136.5

142

The longest scheduled mapping interval is 1 hour; however, the time consideredhas been extended beyond this to allow for alignment and memory time before the startof mapping.

In all of the above, it is assumed that a computer program will be written that willenable the computer to command desired changes in attitude based upon the spacecraftposition in orbit, knowledge of the orbit parameters, and time.

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The attitude rate also has a requirement imposed upon it during mapping. It mustbe kept within ±0.02 deg/sec of zero on two spacecraft axes, and of 3 deg/min on the thirdaxis. The ECU'S supply attitude information to the computer in the form of pulses witha scale factor of 39. 55 arc-sec/pulse. The computer can detect the spacing betweenpulses to better than 10 ms. A rate of 0. 02deg/sec corresponds to 1. 82pulses/sec, or aspacing of 0. 55 second between consecutive pulses. Therefore, the computer can de-tect the maximum desired rate about zero on two axes to within 2 percent. The maxi-mum allowable rate of 4. 2 deg/min about the third axis produces a spacing of 0. 156 sec-onds between consecutive pulses, and therefore the computer can detect the maximumdesired rate about 3 deg/min to within 6. 5 percent. It is assumed that the computer canbe programmed to issue correction commands when the desired rate inputs are exceeded.

.BLOCK H RELIABILITY VERSUS TIME

The Apollo Block II G & N System (with no modifications, spares, or redundancy),when applied to the lunar polar orbit mapping mission, will yield an overall successprobability through entry of .91112. This number was arrived at by combining modulereliabilities into major subsystem reliabilities, and combining these, in turn, to yieldthe overall G & N System reliability. These reliability values are listed below, and thesystem reliability versus mission time is shown in Figure 23.

Subassembly

IMU

PSA

ECU

Optics

D f c C

Total G & N

Probability of Success

.96962

.97648

.96387

.99886

.99955

.91112

Module excitation times were determined from the times assigned to complete the^various functions in the time line. "K" factors, as defined by NAA, have been appliedco the equipment operating times to account for more severe environments encounteredduring the thrusting phases.

Module failure rates, where applicable, were based upon the Block I modulestressed failure rates defined in MIT Report R-429, "Reliability and Quality AssuranceProgress Report," dated December 1963. Some module failure rates have been updatedfor Block H, where equipment design has solidified. In both of these areas, the stressed

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TtUi document oontatn* laloftaai^mtifecti/ns the natlontl dafena* of the United SUtoa withintht meanlnj of the ••plonmge UwaT^&^ljU. S. C. Section* 7M tod 704, the tru»ml«alonor revelation of which In any a^mor tonrhmtuthorlied pereon !• prohibited by law.

SPARK PlUG-th. Elect^nlts Divljlon of GENERAL MOTORS CORPORATION

failure rates were developed from the basic part failure rates as modified by a stressfactor based on the temperature environment and the ratio of operating to rated power _or voltage. In the case of the electronic coupling units, only a parts count and anaverage failure rate have been used without any attention being paid to circuit applica-tions since these units are still in the design development stage. The failure rates •_,used for all parts and modules are found in the Appendix. j

Employing these usage times and failure rates, module reliabilities were calculated -by assuming an exponential distribution of times to failure. Certain equipment reliabilities [were omitted from the calculations because they are not used in the defined mission, orthey are not considered essential and the mission can be adequately performed without n

them.* Therefore, their failure could not cause a basic G & N System failure. All ; jother module reliabilities were combined as though all modules were connected serially,and that any module failure would produce a G & N failure in the mission. A failure ^effects analysis would also eliminate certain equipment from the reliability calculations; , jhowever, this was not done because it was found that the results of a "quick look"failure effects analysis are useless. n

*The following equipment was deemed nonessential:

IMU Temperature Alarm Module, ~\Optics Two-Speed Switch, ' jTracker X Channel,Tracker Y Channel nModulator and Loop Compensation <Star PresenceCosecant Amplifier ^Photometer Electronics (PSA) '• jG & N Panel Brightness ControlOptics Head Tracker Equipment qjOptics Head Photometer Equipment . \Radar Mode ModuleSextant Trunnion 2x ResolverSextant Shaft Computing ResolverSextant Shaft l/2x ResolverTrunnion Angle Counter r~]Shaft Angle Counter jMap and Data Viewer

n

-54- SID 64-1860-14

This document contain* InformattonaffecUpKthe national defense of the United State* withinthe moaning of the eiplonage law'<%j&HE^LS'C' SectlonB 793 "d ?M. the tranimlailonor revelation of which In any ougjjwr tff^^unauthorized peraon u prohibited by law.

SPARK PLUG-the Electronics Division of GENERAL MOTORS CORPORATION

PROBLEM AREAS

PERFORMANCE

There are no performance problem areas. The G & N System is expected toadequately meet all the performance requirements of lunar orbit navigation and localvertical stabilization.

RELIABILITY

As shown in the analysis, the reliability apportioned to the G & N System in the lunarpolar orbit mission (.9975) is not met. Also, it is shown that the inertial measurementunit (IMU), the power and servo assembly (PSA), and the electronic coupling units (ECU)are each below the required overall reliability at the end of the mission, while the sex-tant and scanning telescope head assembly reliability is marginally above the overallG & N requirement.

Each of these subassemblies contains certain low-reliability or life-limited piecesof equipment to which their overall low reliabilities can be attributed. The troublesomeelements are:

Inertial Measurement Unit

Slip ringsBearingsResolversTorque motorsPIP suspensions

Power and Servo Assembly

Gimbal servo amplifiersFail indicator module3200 cps, 1 percent amplifier800 cps, 5 percent amplifier3200 cps AAC filter800 cps AAC filter800 cps, 1 percent amplifier

IRIG suspensionsIRIG and preamplifier assembliesPIPPIPA preamplifiers

-28vdc power supplyPIPA calibration modulePulse torque power supplyBinary current switchDC differential amplifier and

precision voltage reference

In the ECU, all modules have too low a reliability factor except possibly the 4-voltpower supply and the clock and mode logic modules,which are marginal. The opticssubsystem failure rate is due mainly to bearings, resolvers, and motor-tachometers.

-55- SID 64-1860-14

This documentth0 tHt»»nll*f of the

or revelation of which In

I the national detenu of the United Bute* within£.S.C. Sectlona 793 and 794. the tnumliilon

ier to an unauthorized penon It prohibited by law.

SPARK PLUG-tht Electronics Division of GENERAL MOTORS CORPORATION

From the above, it is seen that the types of modules that require reliability im-provement are electronic modules and life-limited mechanical or electromechanicalmodules.

The designs on most Apollo Block II electronic modules have not been optimizedas yet. This is especially true on the ECU'S. Effort has just begun to determine stressfactors, based on circuit applications, to update the module failure rates and to deter-mine where the design may be changed to improve performance and reliability. Pre-liminary indications show that the ECU reliability is better than the figures obtainedon a parts count and average failure rate basis. These investigations will be continuingfor some time, and therefore it cannot be predicted exactly which modules can be im-proved and by how much.

n

nThe problem areas in the mechanical and electromechanical life-limited equipment

all seem to be due to slip rings, bearings, and brushes. Slip rings tend to becomenoisy due to sliding contact wear. This wear is a function of the materials used, thequality of fabrication, temperature, and lubrication. Also, to a lesser extent, con-tamination, either external or caused by chemical action of the parts themselves, cancause degradation of the slip rings.

Brushes appear to be affected in general by the same causes. In addition, brushwear is approximately proportional to shaft speed, and therefore particular applica-tions become important.

On the other hand, bearing failures are due mainly to lubrication problems be-cause of size, temperature, and humidity restrictions. They are also affected byfabrication quality and application.

n

nn

-56- SID 64-1860-14

Thin dominion! contain* lnformaU» nwtnlni o» Uw »plcma(n lawn.or r«v0Utlon at which In Any m

<" nulloiwl 'Infon.o at tho United 8UU> withinT. Norllon. 70,1 ud 704. tho IruiimlMlon

ffler to »n"lmiiuthorUe(t perlan (• prdttbttad by Uw.

SPARK PlUG-thi Electronics Division of CENERAL MOTORS CORPORATION

LIFE EXTENSION ANALYSIS

Several approaches have been investigated to bring the G & N System reliability upto the required level. System reliability can be increased by improvement of equip-ment, modification of equipment usage, and the addition of spare and/or redundantequipment. It should not be overlooked, however, that once the complete Apollo Block nSystem has been defined, a detailed failure effects analysis can be performed and allmodules can be assigned stressed failure rates based upon their component applications.This will certainly improve the accuracy of the reliability predicted for the basicApollo Block II G & N System.

To complete this study, life extension will be discussed oh the basis of the presentreliability predictions.

AREAS OF IMPROVEMENT

Modules

Since the design of much of the Block II electronic equipment has not been com-pleted or optimized, it is not possible to define those areas where modules have beenimproved. Similarly, performance of Block n mechanical and electromechanicalmodules is still being defined, so here also it is not possible to say which modules canbe improved and to what degree. Life tests under various environments are continuouslybeing conducted on all Block n equipment in order to determine life limiting causes andmethods of improvement.

Alternate G & N Approaches

To reduce certain module usage times, and thereby improve their predicted re-liabilities, power for inertial component suspension and for the PIPA loops can beswitched off until just prior to their use. Previously, suspension power was appliedthroughout the entire mission and the PIPA loops would operate whenever the ISS wasoperating.

Using the mechanization shown in Figure 24 to switch on the suspension power15 minutes before ISS turn-on, the usage time is cut to one-fourth of the original onthe following modules.

3200 cps AAC filter3200 cps, 1 percent amplifierDucbsyn transformerPIP suspensionIRIG suspension

_57_ SID 64-1860-14

ThU document contains liifo rmstRn>ajfectlngJh<?patlon«l defense of the United States withinth* meaning of the espionage UwjyJJtfeagr^S1.1C. Sections 793 and 794, the transmissionor revelation of which In any tfSnner to anSftiuthorlied person Is prohibited by law.

SPARK PLUG-the ElecttonkiJlvWon Of GENERAL MOTORS CORPORATION

28vdc Standby Power28vdc to Suspension Power Supply

SuspensionOff

Lamp

Figure 24. Mechanization to Allow Suspension Turn-Off

Switch

T 28 vdc to PIPA Loops

28vdc Operate Power

FadingResistor

-28 vdc Jo we r

+32 vdc Power

+120 vdc Power

28 vdc Operate Power

PIPAOff

Lamp

ExtraPIP

Heater

-28 vdc for PIPA Loops

+32 vdc for PIPA Loops

+120 vdc for PIPA LOOPS

28 vdc to 120 vdcPower Supply

+120 vdc for Gyro* Pulse Torqulng

28 vdc to GyroPulse Torqulng

O—-K Existingjji Fine

AlignRclny

n

n

."-I

Figure 25. Mechanization to Allow PIPA Loop Turn-Off

-58- SID 64-1860-14

This document rnntnlnr l|tffllir"tt"' tdf*""f the national defense of the United State* withinthe meaning of the «imiQnagej!Wfc 11''"^f,-,ff-fi-c Sections 793 and 794, the transmit!ionor revelation of which I$£t5ty manner to an unauthorized peraon ta prohibited by law.

SPARK PlUG-th* Electronic* Division of GENERAL MOTORS CORPORATION

Using the mechanization shown in Figure 25 to switch on the PIPA loop excitation5 minutes before thrusting, the usage time on the following modules can be radicallyreduced.

PIPA calibrationAC differential amplifier and interrogatorBinary current switchDC differential amplifier and precision voltage referencePIPPIP preamplifierPulse torquing power supply

With the above modifications, the G & N probability of success in the lunar mappingmission is . 936. The new subassembly reliability breakdown is as follows.

Subassembly

IMU

PSA

ECU

Optics

D &C

Total G & N

Probability of Success

.982

.9905

.96387

.99886

.99955

.936

At present, the reliability effects due to switching power on and off are 'not known.For the above modifications, PIPA loop switching has been reduced while that for thesuspension was increased.

There are a limited number of built-in alternates for performing the requiredG & N functions as shown in the Appendix. In many instances, the same modules areused, but in alternate modes. A detailed failure effects analysis would indicate whetheran alternate approach is available after a module failure has occurred and what theresultant improvement in reliability would be.

SPARES/REDUNDANCY

Based upon G & N System reliability after modifications, the use of spares and/orredundancy was investigated to increase the overall success probability to the requiredvalue of . 9975 for Profile 1 (lunar polar orbit). The effects of sparing or using redun-dancy were investigated at the module, major assembly, and system levels. Based

-59- SID 64-1860-14

Thle document contain* Infothe m»«nlrn of the eeplonigo 1«*or revelttlon of which In uiy

nil dofonee of the United BUUe withinrs.C. Suction* 7M and 7M, the truiemleilon

rlEed person It prohibited by Uw.

SPARK PlUG-the Electronics Division of GENERAL MOTORS CORPORATION

upon the reliability gained, the additional weight penalty, the malfunction isolationcapability, packaging, the effects on G & N accuracy, the feasibility of sparing underspace conditions, and time periods when sparing is not possible, a suggested combina-tion of modules, assemblies, spares, and redundancy was developed. These are asfollows:

• Redundant heater power slip rings in the IMU,

• One spare IMU and spare inertial component calibration modules in the PSA(44.6 pounds),

• Five redundant ECU'S, one redundant 4-volt power supply and clock and modemodules (23 pounds),

• One spare D and C panel (13 pounds),

• One spare each of the following PSA modules (10. 4 pounds)

1

Gimbal servo and align amplifierFail indicator and warning3200 cps AAC filter3200 cps, 1 percent amplifier•-28 vdc power supplySextant motor drive amplifierScanning telescope motor drive

amplifier800 cps optics compensationServo integrate relay

Anti-creep circuitScanning telescope trunnion trans-

former and relayTracker mode relayScanning telescope mode relayG & N subsystem filter800 cps AAC filter800cps, 1 percent amplifier800 cps, 5 percent amplifier

It is seen that 91 pounds of additional equipment, consisting of 68 pounds of sparesand 23 pounds of redundancy, are needed. With these, the G & N probability of successin the lunar polar orbit mapping mission was calculated to be . 997512.

nn

Subassembly Probability of Success

.999445

.9996

.999622

.99886

.99983

.997512

-60-

This document contain! Infor:the meaning of the espionage lawa,or revelation of which In any

SID 64-1860-14

ma] defense of the United State* withinJTS.C. Sections 793 and 794. the truumlMlon

lauthorlred per»on ii prohibited by l»w.

SPARK PlUG-tht Electronics Division of GENERAL MOTORS CORPORATION

In all the above calculations, it was assumed that the opticlTfiead could not be sparedor made redundant since this would constitute a major G & N or spacecraft redesign.With this constraint, the weight and volume penalties become critical if more of theother equipment is added to increase the overall G & N success probability.

Some form of redundancy was required in the case of ECU's to improve theiroverall reliability during critical periods. Critical periods when sparing could not beperformed are: entry, all thrusting periods, 1 hour prior to deboost to lunar orbit,and 1 hour prior to entry. It was considered that all other thrusting periods could bepostponed if sparing was necessary. For the reliability calculations, each of the fiveECU's and the 4-volt power supply and clock and mode logic modules were consideredto be paralleled by identical units. An alternate approach would be to design redundantparts into the existing modules.

A possible implementation for paralleling entire coupling units is shown in Figure 26.The redundant ECU operates independently and in parallel with the primary unit, withthe outputs controlled by switches. Normally, the outputs of the primary ECU wouldbe used and those of the redundant ECU inhibited by the output switches. The switchingfunction for the output switches is provided by the failure detector, which monitors sixselected signals. It is conservatively estimated that 90 percent of all possible failuresin the ECU would be detected by the monitoring of these six signals. If a failure shouldoccur in the primary system, the signal outputs used by other subsystems wouldbe switched from the primary ECU to the redundant ECU.

The six primary ECU signals selected for monitoring are: +14vdc, +4vdc, 800 cpsreference, digital-to-analog converter failure, fine error, and coarse error. The+14vdc would be monitored in the analog mode module and would serve as a test forthe operation of the 14vdc supply, as well as the 28vdc source, which the 14v supplyuses for primary power. The +4vdc supply would be monitored at the error anglecounter module, and would serve as a check on digital logic power. Both the 14v and4v signals could be monitored by a device that would give a failure indication (changeof voltage level) if the signal was not within a specified tolerance band. The 800 cyclereference could also be monitored by a similar device after initial detection (rectifica-tion) and integration (filtering) to a dc level. The 800 cycle reference signal would bemonitored in the digital-to-analog converter and coarse error loop. This would checkthe operation of circuits in both modules.

The digital-to-analog converter (DAC) failure signal could conceivably involvecomputer programming. The error counter of the ECU that controls the DAC acceptscomputer pulses as its driving function. A method of checking DAC operation would beto match the solution determined by the DAC to that of the computer, and give a failureindication should a mismatch occur.

-61- SID 64-1860-14

Thli document contain* informatlOTTftCtgcthj^neiiatlonal defenie of the United State* withinth* moaalnf of the espionage Uw§. yriptt&J.8. C. Section* 793 tad 794. the tran*mj»lonor revelation of which In any njpfi£«r to anonauthorlzed penon 1» prohibited by law.

nlcs Division of GENERAL MOTORS CORPORATION

E2bCoJ

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-62- SID 64-1860-14

Thla document conUlu lBformatio?the mMning of the espionage laws.

,.i..,_ „»

5nal defenoe at the United SUte> within" Sections 783 and 794. the tranamlailon

horlred oeraon la nrohlblterf hv imw.

SPARK PLUG-the Electronic. Diviiion of GENERAL MOTORS CORPORATION

• Inertial Measurement Unit

• Total weight is 42 pounds.

• Four captive socket head screws will be used to secure the IMU.

• Three PIPA calibration modules and one IRIG calibration module will be1 matched to the spare IMU, and installed with it.

• The coolant connectors on the IMU will accommodate a connection and dis-connection of the supply hoses in an atmosphere of 7 psia without loss of coolant.

• Replacement of the IMU can be accomplished with the G & N Indicator ControlPanel removed. In addition, the optical shroud may require removal. Somemechanical redesign of panels may be required to minimize G & N disassembly.

• Appropriate grips or handles on the IMU for zero-g handling will be required.

• Power and Servo Assembly

• Average weight of the modules is 0. 68 pound.

• All spared modules are accessable (as presently configured) with only theG & N Indicator Control Panel removed.

• Moisture proofing is accomplished at the module level.

• Captive screws are used to hold the modules to the header and coldplate.

• All receptacles will have guide pins to aid in replacements.

• One size socket head common to all modules.

The onboard storage of the listed spares is assumed to utilize space in the labora-tory Module associated with mission Profile 1. The volume requirements are tabulated,and of special significance is the heat requirement of the IMU. The requirements areas follows.

Inertial Measurement Unit

The volume required to store the spare IMU is approximately 1, 440 cubic inches|(14-inch clearance sphere). The spared IMU, filled with coolant fluid, will require an expansionbellows plug to relieve high pressure differentials in the coolant lines caused by temper-ature extremes. The IMU will also require shock mounting while in storage. This canbe accomplished by providing a shock-mounted plate fitted with mounting pads to whichthe IMU can be mounted in the storage area. Since the IMU will require heat duringthe storage period, a 28 vdc source (approximately 2 watts average) must be availablein the storage area to operate the heaters built into the IMU.

•—^; i

—i

.1

-64- SID 64-1860-14

or revelation of which tn any

SPARK PLUG-tJ Division of GENERAL MOTORS CORPORATION

Power and Servo Assembly

The average volume of a PSA module is 11 cubic inches. Each module will bepackaged separately in a moistureproof bag and will require an outer container forshock and vibration protection. It is anticipated that the average module, with packaging,will require approximately 15 cubic inches. The suggested spares list for Apollo Xcontains 21 PSA modules and will require approximately 315 cubic inches of storagespace.

Display and Control Group

The D & C Group will be packaged in moistureproof bags, and will require an outercontainer for shock and vibration isolation. The D & C Group will require approximately1, 560 cubic inches of storage space.

Malfunction Isolation Considerations

The time required to isolate, remove, and replace a malfunctioned component isbased on an average time of 2.5 hours. It can be seen that (based on the probability offailure for a G & N System with no spares and Malina's Tables of Poisson's ExponentialBinomial Limits as applied to probability of failures) the probability of having more thantwo failures is insignificant. Therefore, the total maximum maintenance downtimefor the total mission will be considered as twice the average downtime for one failure,or 5 hours.

The average downtime was arrived at by sampling probable failures and estimat-ing individual times required to isolate the failure. The 2. 5-hour average downtimeconsists of:

• Data accumulation —1/4 hour

• Communications with earth — 1/4 hour

• Planning and interpretation of data — 1 hour

• Removal and replacement — 1/2 hour

• Verification of repair — 1/2 hour

The fault isolating procedure that will be used is intended to proceed through thefollowing steps.

• Detect failure.

• Switch computers. (If failure clears, computer failure can be checked out withself-check program; if not, system diagnostics can be initiated.)

-65-SID 64-1860-14

SPARK PlUG-the Electronics Division of GENERAL MOTORS CORPORATION

• Switch ECU'S.

• Localize failures to greatest practical extent via downlink. Supplement standarddownlink with system diagnostic programs in the computer.

• If possible, find mode changes, backup methods, or other procedural means forbypassing the problem.

• Determine the failed component via telemetry data,and replace.

• Assist astronaut in manual checkout of the system until failed component isisolated.

These steps are intended to be run in the order listed. At any time, restorationof satisfactory operation will make further progress through the list unnecessary.

The above fault isolation procedure assumes the primary role for isolation of mal-functions on the downlink telemetry with onboard capability as a backup tool. Table 6shows a list of downlink telemetry requirements (based on Block I planning), whichappears to provide minimal needs for isolation to the recommended spares level.Figure 27 illustrates a procedure that can be implemented on a ground computer (orany combination of automatic and manual procedures) to provide the ground malfunctionisolation capability. Implied within this procedure is a variable format telemetryscheme (not presently implemented) that permits expanded sampling and sampling ratesin the computer and PIPA loop areas when a subsystem malfunction occurs. This typeof system has been successfully used on other guidance systems. It essentially per-mits expanded looks at malfunctional areas by temporarily reducing monitoring of otherareas so as not to exceed overall allocated bandwidth. The variable format for thecomputer and PIPA loops could be implemented by uplink command, astronaut operatedswitch, or both.

Because of the extremely integrated nature of the Computer Subsystem with theInertial and Optical Subsystems, diagnostic subroutines in the computer can be im-plemented on a stimuli-response basis to assist in localizing to IMU or PSA electronics.These tests would be called up by astronaut operation of the Display and Keyboard.A ground based program, coupled with system diagnostic routines, can be made adequatefor establishing malfunction isolation, spares, and/or alternate procedures.

The onboard test point availability is expected to be adequate for isolating to thespares listed. If special onboard instrumentation is provided, manual isolation offailures by use of an ac-dc VTVM and an oscilloscope could be accomplished. However,this type of approach would be recommended only in the event of dual failure of G & Nand telemetry. Possible consideration should be given to simplified manual isolationby block replacement of modules (at some penalty in reliability however) to minimizethe specialized knowledge and tedious procedure and intepretation that this type ofapproach applies.

nnr-s. 4

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-66- SID 64-1860-14

SPARK PLU6-lh« El«tronl» Division of GENERAL MOTORS CORPORATION

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-67- SID 64-1860-14

tte ****«"* r of the espionage Uor rtvelatlOD of wblch in

fng the national defence erf the United State* within18, U.8.C. Sections 793 and 7M. the tnuumlaoloo

an unauthorlted peroon 10 prohibited by law.

SPARK PlUC-th. Cfoctronlct Division of 6CNCRAI MOTORS CORPORATION

Program Outputs(Poii., Vel., Tlnii)

:lGeneration of S/C PathOn Ground Computer

MarkTlmea

Pick Out O & N Position atMark Tlnvm —nt H.S. Tlmns

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('ompnro G :tn»l N PnnillonTo nnitar Position ToI'OHltlon from OpIlcH

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Figure 27. Malfunction Isolation Procedures (Sheet 1 of 6)

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-68- SID 64-1860-14

Thlt dooumnt contain lafoitb* -"--1-f at tt» Mplotiai* !«••or rwvUtlcB of wfalob tn any.

national d«t«m. of Ik* Uoltld ftaiM wlthl»Biotlooa 791 and TM. UM trmn.ml.ilon

li*d panon It »rohlblud by law.

SPARK PlUG-th. Eledronlcs Dlvliion of GENERAL MOTORS CORPORATION

Ck All Aliirms

Optlcx

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Figure 27. Malfunction Isolation Procedures (Sheet 2 of 6)

-69- SID 64-1860-14

Tbli document conUtha msulQf of th«or r«TflUtlon of which

'SIng th« national dofona* of th« United StaUi within

[.B.C. 8«otlona 7&3 and 794, the tranimlislotiunauthorlted p«non !• prohibited by law.

SPARK PLUG-Ik Eloctronlci Division of GENERAL MOTORS CORPORATION

iSwitch "on llnu" AGC's

Check Position andVelocity In Second AGC

Call G & N SyHlcm DlvDiagnostic Program

Set Finn A(Cln.ir If AInr inB l.t)

Figure 27. Malfunction Isolation Procedures (Sheet 3 of 6)

Thli document contcJn* Iniotil* mMnlnf of the uplonige lor revelation of which In

-70- SID 64-1860-14

aftectla^nc national defoua tf the United States wlthtn. U. B.C. Section! 783 and 794. the tranemlulon

_ unauthorized peraon If prohibited by law.

SPARK PLUG-tha Electronics Division of GENERAL MOTORS CORPORATION

CDII

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-71- SID 64-1860-14

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SPARK PlUG-th. Electronlci Division of GENERAL MOTORS CORPORATION

n

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to "AGC-MFI" Mode*

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All Program InputsAll Program OutputsAll Interrupts Ind.All Increments Ind.Time of ull of ahoveAll miuling commands (I)SKY)All IN 81 OUT Registers

4

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-72- SID 64-1860-14

Thl> document contmlns InformatloikggcctliiKHie meaning of the eiplonage lawi, Titor revelation of which In any

ftlonil defenie of the United Statee within7s. C. Bectloni 793 ud 794, the trmumliilon

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Division ot GENERAL MOTORS CORPORATION

NOTE: If PIPA spares are available,this loop goes as dhown.

Switch Format To Enable |Hi-H\v Lock At S.G. Output I

If flag is reset, call AGC Self Chuck ,It this is bad, follow UiC procedures,[f not, replace interface modules.

Figure 27. Malfunction Isolation Procedures (Sheet 6 of 6)

-73- SID 64-1860-14

This document contain*th* meaning of the espionageor rwfllatlon of which in

he national defenoe of the United State* withinU.S.C. Sections 793 and 794. th* trtaamiasion" i unauthortied person It prohibited by law.

SPARK PlUC-Un Electronlg Division of GENERAL MOTORS CORPORATION

PARAMETRIC STUDIES

Using the modifications, spares, and redundancy as defined for the lunar polarorbit mapping mission, parametric curves of probability of success versus time areshown in Figures 28, 29, and 30 for Profiles 1, 2, and 3, respectively. Curves ofprobability of success versus weight are shown in Figures 31 and 32 for Profiles 1and 3, respectively. Since Profile 2 does not have a definite operating length, end-point reliabilities could not be calculated for a curve of reliability versus weight.

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SPARK PlUG-th. ElKtroni» Division of GENERAL MOTORS CORPORATION

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SPARK PLUC-tht Dettr^j DMsjo^oLGENERAL MOTORS CORPORATION

RECOMMENDED SYSTEMS.1

CHANGES FROM BLOCK II SYSTEM !

The recommended system at present would consist of the Apollo Block n G & NSystem with the modifications, spares, and redundancy previously discussed. However,dependent upon certain areas where further definition and study is necessary, therecommended spares and redundancy requirements may be lessened.

With a reduction in the overall G & N System reliability requirement to . 9965, orwith further engineering development, the following spares configuration applied to the {modified Block II System appears ultimately feasible.

• One spare IMU and inertial component calibration modules in the PSA j7!!(44. 6 pounds), J '»

• Two spare ECU'S and one spare each of the 4-volt power supply and the clock —and mode logic modules (10 pounds), ^ '•.]

• One spare each of the previously spared PSA modules (10. 4 pounds).F]

The G & N System additional weight requirement would then be reduced to 65 pounds Iof spares and no redundant modules, except possibly some IMU slip rings. The addi-tional packaging and storage requirements would then also be reduced commensurably. R

' . ]

ARE AS REQUIRING FURTHER ENGINEERING DEVELOPMENT ^"i

Stressed Failure Rates and Failure Effects Analysis1 • -n

A significant quantity of the reliability predictions have been based on the average - ' ' ' 'failure rates of the parts and modules. A study of electrical stresses on the individualparts in each module will yield higher reliabilities than previously assumed. Specific HIexamples where minor changes in modules and subassemblies can increase reliability 'will also become apparent. This technique was proven successful in the ECU assembly,where a preliminary parts stress investigation allowed an increased probability of ""1success to be predicted. '<

Once the actual stressed failure rates are determined, a detailed failure effects '• janalysis can be used to separate anticipated failures into: a '

• Major G & N failures for a specific mission, n

• Minor performance degradation.

nFrom past experience, it is known that without a failure effects analysis, failures 1

that result in minor performance degradation are often erroneously predicted as mis-sion failures. _;

, i

-80- SID 64-1860-14

Thil document contain InfofBzmjj^HecUn^he national dafenae of the United State! withinth. meaning at the eiplonage lawi.'lj^^^S. C. Sections 783 and 794. the tnwamlailonor revelation of which In any,jrf>Rr to^SStaauthorlted peraon u prohibited by law.

SPARK PUIG-tJMjgtffliiks DMslon of GENERAL MOTORS CORPORATION

Component and Development Testing

The Apollo G & N equipment approaches the state of the art in the areas of com-ponent usage and safety factor design. However, some of the remaining failure causeshave been traced only to major categories. This is particularly true of slip ringsbearings (gimbal, motor-tachometer, and gyro), and torque motor brushes. Testingunder conditions that simulate actual usage should be conducted on these componentsto obtain specific examples of failure. In this manner unique failure modes peculiaronly to the G & N equipment can be identified.

Switching Stresses

Power turn-on, turn-off, and switching effects on failure rates and modes of theG & N equipment have not been fully investigated. It is recommended that this be done.

Optics Equipment Improvements

The optics head had a probability of success that barely exceeded the G & N designgoal. Therefore the remaining portions of the G & N equipment would have had to achieveprobabilities of success approaching unity. A detailed investigation of the componentsof the optics head should therefore be conducted. If this investigation is followed bya more accurate trade-off and apportionment study for the remainder of the G & Nequipment, there is a good chance of eliminating some of the on-board spares presentlyrecommended.

Alternate Procedures

An engineering investigation of the procedures used to perform G & N missionfunctions is needed. One such investigation resulted in a new procedure for alignmentof the inertial reference, which reduced the operating time and simplified the effortrequired of the astronaut. Another preliminary study on the failure of components usedin the optics head revealed that alternate procedures can be established which, withsome loss in convenience and accuracy, will allow the astronaut to perform the G & Nfunctions necessary to mission success. These investigations should be continued andexpanded into other G & N areas. Alternate procedures, to be used in the event ofcomponent failures within the G & N systems, should be established to increase crewsafety and the probability of mission success.

Redesign Summary

It is believed that the majority of G & N system components are "state of the art-"However, redesign may be worthwhile in the light of technological advances and thestudies of stressed failure rates, failure effects, and alternate procedure outlinedabove.

~81~ SID 64-1860-14

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c. SPARK PLUG-On Ekctronlct Dlviilon of GENIRAl MOTORS CORPORATION

IV. TEST PLANS

INTRODUCTION

The purpose of these test plans for the Apollo X Guidance and Navigation (G & N)equipment is to assure that the G & N is developed, manufactured, and tested as aqualified subsystem for the Apollo X mission. The sequence of the Apollo X programprogression is shown in Figure 33. The proposed sequence is intended to yield thefollowing advantages:

• Design goals and achievements are subjected to continuous scrutiny with respectto critical mission requirements,

• Manufacturing and inspection integrity are continuously evaluated,

• Potential problem areas are screened several times to eliminate human error,

• Trouble spots are exposed early in the sequence when corrective action canmost judiciously be initiated,

• Some measure of reliability and predicted performance has been establishedbefore the program of flight tested reliability is started,

• Additional information is available for establishing reliable operation andmaintenance requirements.

The G & N System development schedule is integrated with the Lunar Apollo Pro-gram (see pages 114 and 115) and represents the best estimate to date of:

• Least conflict of critical facility usage,

• Maximum utilization of present and projected field operations,

• Minimum necessary research and development activity,

• Maximum usage of equipment qualified by the Lunar Apollo Program.

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SPARK PLUG-tha Electronics Division of GENERAL MOTORS CORPORATION

SCOPE

The planning for this study will be presented in four major test areas:

« Design and development,

• Acceptance and delivery,

• Qualification,

• Interchangeability.

Each of the areas will be expanded upon separately and recommendations madefor Apollo X. Because of the similarity of Apollo X to the Lunar Apollo equipmentand mission, the Lunar Apollo will be considered as a base guideline. Therefore,only additional or unique requirements for Apollo X testing will be fully investigated.The additional constraint is that no planning has been considered for formal reliabilitydemonstrations in compliance with a NAA request.

SUMMARY

All reliability assessments, quality assurance, manufacturing techniques, hard-ware and design change control, program control and review, liaison procedures,field support, and cost accounting methods presently in existence on the Lunar ApolloProgram are recommended to be adapted to the Apollo X Program.

The modifications proposed for Apollo X and the life limited equipment will under-go a reliability and research and development testing program, starting with the defi-nition phase of the project. This program will continue until no further modificationsto the G & N can be justifiably proposed. Wherever possible, the results of benchtest programs of the Lunar Apollo will be utilized.

An interchangeability and design review testing program will be conducted with a.G & N subsystem to evaluate the effects on G & N parameters and performance of on-board spares, repair, and redundancy.

It is recommended that no testing should be done on the electromagnetic inter-ference and susceptibility of the Apollo X G & N equipment. Results from the LunnrApollo Program can be considered as typical performance of the Apollo X G & N.

The Apollo X G & N is considered qualified for mission vibration, acceleration,and shock by the Lunar Apollo qualification program and normal manufacturing controls,with the exception of the Apollo X PSA package, redundant ECU package, modifiedD & C group, and on board spare IMU. These units will be qualified for dynamic en-vironments.

-85- SID 64-1860-14

ThlB document contains inftffiiB&lon ailflOTfng the national defenae of the United Statea withinthe meaning at the espionage lawftJomelS, V.S.C. Sections 793 and 794. the tranamlealonor revelation of which in in^manne^to, an unauthorized peraon Is prohibited by law.

SPARK PUIC-thtJbctrMla Division of GENERAL MOTORS CORPORATION

One Apollo X G & N wiH~bT^ubjected to the temperature-altitude-humidity testingrequirements of ND 1002037, Apollo Airborne Guidance and Navigation Equipment En-vironmental Qualification Specification, with the Inclusion of the 50 percent nitrogen and50 percent oxygen atmosphere.

One of each type of Apollo X modules will be exposed to the two-gas atmosphereunder temperature-altitude conditions.

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Thl» document contain* 'th» mcanlaf of the etplon.go Uor r«relatlon of which In any

national defen» of the United SUtu within. C. Section! 7M ud 794. th. tranimllilonIiuthorlied perion I. prohlbltnl by l«w.

~ SPARK PLUG-the Electronics Division of GENERAL MOTORS CORPORATION

TEST PLANNING

DESIGN AND DEVELOPMENT TESTING

General

During the program definition phase of the Apollo X Program at AC Spark Plug, amilestone will be reached at which time an integrated design and development benchtest program will be justified, initiated, and supported. Sufficient detail informationwill be available to preschedule specific development test tasks between AC Spark Plug,Kollsman, and Raytheon.

All design and development testing considered herein presupposes that in somespecific instances:

• Apollo X requirements are not the same as the G & N Block n,

• Lunar Apollo requirements are modified and anticipated similarity of dualfunctions cease to exist,

• To proceed on a noninterference basis with Lunar Apollo schedules parallelefforts may be required.

Design and Development Testing Objectives

The objectives of the design and development activities and testing for the Apollo XG & N System are to assure the following.

• Preliminary designs, mechanizations, and goals established by the study programare optimized and integrated with a realizable equipment and testing schedule.

• Reliability of the G & N spacecraft equipment is maintained, improved, anddefinitized throughout the program.

• Complete manufacturing definitions, test plans, test requirements, and adequatesupport equipment are specified, evaluated, proof tested, and supported through-out the program. This includes both inhouse, subcontractor, and field locations.

• Adequate electrical and mechanical G & N interfaces with the spacecraft are deter-mined and maintained through sufficient liaison with NASA and NAA.

• Necessary spacecraft crew utilization, maintenance, and emergency repaircapabilities have been reflected into the design of the G & N equipment.

_8 7_ SID G4-1860-14

Thia document contain* InformaT&te&ectlBl&Oie national defense of the United State* withinthe mining of the eaplonage lawa, 7jjM»?S»83 .£SJSectIcm8 793 and 794, the tranamleelonor revelation of which In any majSfr to an unauthorized person tg prohibited by law.

SPARK PLUG-the Electronics Dlviilon of GENERAL MOTORS CORPORATION

Design and Development Testing Plan Implementation

During the definition phase of the program, several areas of design improvementsand test results required to obtain increased reliability will have been investigated andchanges defined or anticipated. These are expected to include requirements for: ~?

• Redesigned PSA module packaging for flight replacement and maintainability,

• Unique packaging and test requirements of redundant ECU'S, 1

• Mechanizations, circuits, and procedures for accelerometer loop turn-off andinertial component suspension turn-off, -^

I

• Redundant slip ring usage where indicated by reliability studies for temperature •control, attitude control, PIP A loops, and IMU power interfaces,

• Minimization of vibration transmissibilities to critically aligned components :such as navigation base — IMU — optics alignments and inertial components,

• Reduction of electrical stress factors within modules by partial component re- idundancy and upgraded components,

• Mechanization simplifications for the reduction of possible failures and malfunc- r"-tions, {

• Specific recommendations for engineering test evaluations of life-limited rotating _components, environmentally affected sealing, and pressurization materials and | ,!components,

• Integrated proof testing of Apollo X G & N equipment change impact on support ~~and handling equipment and associated test procedures,

• Verification of any new guidance computer flight, ground test, field test, and —,inhouse test routines, ;i

• Establishment of initial electrical and mechanical spacecraft interfaces,

• Integrated task analysis with respect to inflight maintenance and onboard sparesreplacement,

• Confidence checks of failure effects analysis, error analyses, and test tolerance ~studies,

• Thermal evaluation of proposed mechanizations. —i; J

All of the above anticipated developments will generate logical demands for initialcircuit breadboards and testing, both for flight and ground support equipment. These r-»should be followed up with engineering developmental proof models as warranted by the i-|particular test results and analysis. In addition, ground support equipment andApollo X G & N flight hardware compatibility should reasonably be expected to be estab- —\lished and maintained. To adequately define spacecraft mounting configurations, amechanical gage or mockup Apollo X G & N model would be required.

_88_ SID 64-1860-14

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SPARK PLUG-the Electronics Division of GENERAL MOTORS CORPORATION

Design and Development Data Requirements

All test data from the Lunar Apollo Program will be reviewed. In addition, cur-rent results from the Apollo X testing will be reviewed. Reliability assessments ofmodules, major units, and components must be available and augmented by failurerates unique to Apollo X G & N equipment. A moderate amount of digital computermachine time will be required for servo loop analysis and new guidance computerroutine operation. Adequate analog computer time will be used to accomplish the de-sign objectives.

Design and Development Equipment Required

To support the design and development objectives will require:

• One G & N Block II System, updated with developmental proof models to anApollo X configuration. This includes requirements for developmental proofmodels of the Kollsman optics and the Raytheon computer withtwo display keyboards.

• One G & N Block II GSE and universal test site support and calibration equip-ment, updated to Apollo X configuration, with developmental proof models.

• Flight hardware and GSE breadboards.

• One accelerometer *\ „. . , , , . . , . .V Status and evaluation test area• One gyroscope J To be determined at time of receipt

• Miscellaneous breadboard parts.

• One mechanical gage system.

Design and Development Facilities

Approximately 30 x 30 feet of floor test space with a 13. 5 x 13. 5 foot vibrationisolated pad is required. In addition, approximately 15 x 15 feet will be required forthe mechanical gage system, and approximately 20 x 20 feet of circuit breadboard areawill be required Usage of all areas will be integrated with Lunar Apollo where possibleon a noninterference basis

ACCEPTANCE AND DELIVERY TESTING

Acceptance and Delivery Testing Objective

The objective of acceptance and delivery testing is to maintain manufacturing andinspection integrity with respect to material, component, module, subassembly, sub-system, and G & N performance.

_8,,_ SID 04-18(50-14

Thla document contain* informatlornBtartlnstha^atlonal defence of the United States withinthe meaning of the espionage laws, TIUe^JwTs.c- Sectlone 793 and 794, the transmissionor revelation of which In any trtjSaerto an^&authorlted person le prohibited by law.

SPARK PlUG-«h. Electronics Division of GENERAL MOTORS CORPORATION

Acceptance and Delivery Testing Plan Implementation

To accomplish the testing objective, the test flowgram of Figure 34 is proposedfor the Apollo X program. The flowgram sequence is very similar to the type employedon the Lunar Apollo flowgram. The main areas of difference are: an additional spareIMU will be tested, a PSA-ECU combination with redundant ECU'S, a minimallymodified G & N Indicator panel, and the additional onboard spare subassemblies andmodules testing.

Testing of the spare IMU will be accomplished by: serial testing of a G & N withtwo IMU's, parallel testing two ISS configurations, or by the use of a standard ISS con-figuration for testing all spare IMU's. The method of testing will be determined by theintegrated schedule of equipment usage. A preliminary study indicates that a standardISS configuration for testing all spare IMU's represents a minimal equipment usage.The onboard spares can be tested at the subassembly or module level; however, theinterchangeability testing program will verify the assumption. Should the interchange-ability program prove this assumption incorrect to any extent, the integrated testingof spares at the ISS, OSS, or G & N level can be integrated without any additionalfacilities or test equipment. .

Checkout and initial testing of the optics and map and data viewer is proposed tobe done at Kollsman Instrument Corporation. Checkout and initial testing of the guid-ance computer and display and keyboard is proposed to be done at Raytheon. However,AC Spark Plug will perform receiving inspection tests on these components prior toacceptance.

The following areas of activity will be necessary, starting with the program defini-tion phase and continuing through to the end of the delivery program to assure timelydelivery.

• Initiation and continuing monitor of PERT planning,

• Documentation of manufacturing effort, that is, manufacturing progress anddocuments defining equipment and building techniques,

• Procurement of material, components, and vendor selection,

• Monitor vendor and subcontractor scheduling with respect to delivery schedule,

• Determination of plant layout and integrated usage of manufacturing facilities,

• Determination, design, evaluation, procurement, maintenance, and calibrationof test equipment,

• Development and implementation of fabrication and assembly techniques,

• Monitor and implementation of adequate manufacturing personnel training,

_90_ SID 64-1860-14

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SPARK PlUG-the Electronics Division of GENERAL MOTORS CORPORATION

IMU

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-91- • SID 64-1860-14

This document contain Inforthe meaning of the eaplonage lawsor revelation of which In any

(e national defense of the United States within!.C. Sections 793 and 794, the transmission^ ""lorlzed person Is prohibited by law.

SPARK PLU6-th« Electron^ Division of GENERAL MOTORS CORPORATION

• Institution of corrective action and resolution of manufacturing and testingproblems, ~~j

' i• Assistance and liaison with Engineering during initial breadboard phase,

• Determination of adequate lead times for critical components, subassemblies, ~jand processes for timely manufacturing and delivery, 1

• Determination and maintenance of adequate quality assurance and test records. ~,\i

The activities previously discussed will rely heavily on experience gained fromthe Lunar Apollo Program, and will employ the same techniques. ~-

!

Data Handling

^Complete data history and lot control of incoming materials, parts, modules, sub- .' ! jassemblies, and major units will be maintained. These records will contain status, ilocation, and disposition of all articles. In addition, all testing data required by FTM, "~j jATP, and JDC must adequately be monitored, prepared, reproduced, distributed, and • ' !kept on file for evaluation. PERT planning and summary sheets will periodically be _^ '.required. Failure reporting and corrective action summaries must be periodicallydistributed.

i—i iEquipment Required <

The acceptance and testing program will require a complete GSE to equip a univer- Hsal G & N ISS-OSS test site, a complete GSE to equip an ISS-OSS test site, and a com- ;

plete set of GFP, GSE, and GOE auxiliary equipment, and site equipment to supportboth test sites. This equipment's integrated usage with the Lunar Apollo Program is "to be determined. . ' •>

A 16 PIPA test console (GFP) and support equipment will be required, which includes ~~]as major items a precision rotary head (GFP) and a 36 x 36 inch leveled surface • Jplate. In addition, modification kits for present inhouse test fixtures will be updatedfor Apollo X G & N equipment. The integrated usage of this equipment with the Lunar HjApollo Program is also to be determined. ,'

—*•*

Acceptance and Delivery Testing Facilities Requirement j

To support the acceptance and delivery testing, a minimum of 50 x 50 feet of floor r-jspace of test area is required. This test area should contain a 13. 5 x 13. 5 foot isolated ; } ;pad and four precision aligned optical targets. In addition, the environment of the test ' :

area should be as follows. —,

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• Air conditioned (temperature 75° ± 4°F, relative humidity 50 percent or less)r"if i • Pressure differential (internal positive pressure of 0.05" H2O)\ \

• Filtering method (final filter HEPA type)

j i • Particle count tolerance (determined by activity, size, and product throughempirical analysis)

p • Particle size and count (adjacent to optical unit). A count of 100,000 particlesI per cubic foot of 5. 0 micron or larger. This requirement applies only to periods

when the sextant and scanning telescope optics heads are uncovered.

I • Hoods (when required to contain dust sources or vapors)

• Volatile vapors (toxic vapors vented through hoods and air conditioning)

• Floors (smooth waxed tile)

• Interior finish (hard finish paint)

• Utilities and fixtures (standard benches ~ open storage)

• Cleanliness (layout with consideration to cleaning access —cleaning weekly ormore frequently, as required)

• Cover garments (caps, gowns, and boots)

• Entrance chamber (required)

• Additional floor space is needed to support inertial component prealignment

INTERCHANGEABILITY TESTING

The Apollo X is predicated on onboard flight-replaceable modules, subassemblies,and components of the G & N System. A formal program of interchangeability testingis an important part of an integrated test plan to demonstrate the integrity of G & Nspacecraft interfaces when spares are inserted in the G & N. The need for tested inter-changeability is a unique requirement for Apollo X. No similar-type test is conductedon the Lunar Apollo Program, although the design itself reflects interchangeability tothe presently scheduled field sparing levels.

Interchangeability Testing Objectives

The objectives for interchangeability testing are as follows.

• A formal demonstration that modules, subassemblies, and major portions canbe replaced in the G & N without recalibration and without degrading the per-formance.

• To establish the relationship between G & N parameters and spares replacements.

-93- SID 64-1860-14

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SPARK PLUC-(h« Electronic! Division of GENERAL MOTORS CORPORATION

Interchangeability Testing Plan Implementation1

1Intel-changeability testing will be conducted at AC Spark Plug by Manufacturingand monitored by Quality Assurance. Initial data analyses will be the jointresponsibility of Quality Assurance and Engineering; final data analyses will be the ~~*responsibility of Engineering. The G & N will be installed in a universal test station 1area capable of conducting tests with both ISS-OSS and G & N configurations. Afteran initial G & N acceptance and delivery test to establish a reference for further data ~"!evaluation, each module, subassembly, or major component of the onboard spares »will be substituted into the G & N. Testing will then be conducted to evaluate thechanges in G & N parameters. The process will then be repeated until every piece of ""]the spares has been evaluated in the G & N. Where necessary, the G & N will be 'rearranged into an Inertial Subsystem or Optical Subsystem configuration to accomplishthe evaluation of the spare replacements. This interchange ability test flowgram is shown ~~]in Figure 35. All testing will be conducted at normal room ambient environmental con- . <ditions. Upon completion of the interchangeability testing, necessary refurbishmentand formal testing can be accomplished on the G & N and used to update or support as \a spare the NAA environmental test vehicle, ground test house spacecraft, or the NAA - •G & N Block n house spacecraft. Except for the refurbishment, this disposition is _^most feasibly accomplished on site. <

Proof of Accomplishment .—i]

The satisfactory completion of those tests in the final test method (FTM) that beara significant relationship to the module, subassembly, or spare unit replaced in the ~|G & N will constitute an adequate demonstration of interchangeability. The tests con- iducted will be of the following, but not necessarily limited to the list.

• Standby Power-On Test ; ;

• Operate Power-On Test

• Failure Indicating Circuitry Test

• IMU Temperature Control

• G & N System Power Supply Testi

• Map and Data Viewer Test

• G & N Panel Brightness and Lamp Test j

• Guidance Computer Operational Test

• Zero Optics Test !j

• Optics Slew Rate Testn

-94- SID 64-1860-14

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SPARK PLUG-Ik Electronics Division of GENERAL MOTORS CORPORATION

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SPARK PlUG-tht Electronic* Division of GENERAL MOTORS CORPORATION

• Optics Resolution Test

• Guidance Computer Mode Control Test ~j

• Manual Mode Control Test

• Minimum Impulse Controller Test '"]

• IRIG Scale Factor and Gimbal Torque Test

• Gyrocompassing Test i

• IRIG Coefficient Determination and Simulated Space Fine Align

• PIPA Scale Factor Test !\The final data evaluation will determine the completion and adequacy of the

demonstration. :"""].• i

Interchangeability Testing Data Requirements -^t

Data Acquired!

Data to be analyzed and reviewed will be specified by the particular job descriptioncards ( JDC) performed. This will constitute the main bulk of data. Additional datarequired will be: ~)

\i

• Gyro, accelerometer, and optics characteristics at periodic intervals,

• Inertial components temperature log, <

• G & N System test environments and operating time,

• G & N test configuration,t

• Sufficient information of subsystem, system, and assembly serial numbers toenable accurate time correlation techniques. •-*

Data Analysis _^

The data analysis of the interchangeability testing can be divided into three maincategories: JDC required analysis, Engineering analysis and justifications, and com- —puter parameter variations. JDC required analysis, essentially a "go," "no-go" basis, iwill be used as a criterion for proceeding with the testing. The final data evaluationwill use any combination of correlation and regression, statistical variance, difference — \equations, or analog simulation to effectively analyze the signal or parameter variations. jThla type of data analysis will be on an Engineering basis. Analysis and justification

_96_ SID 64-1860-14

Thle document contain* information attQCttng the naUGnal defense of the United State* wlthlflthe -—-i-g of the eeplonage lawi. TltleltMU^T Section! 793 and 794. the tranamlMlonor revelation of which In any manner to^tir'ahx^horli.d perion le prohibited by law.

SPARK PLUG-th. Electronlci Diviiion of GENERAL MOTORS CORPORATION

will also be performed for any necessary adjustments which may have to be made tomaintain G & N tolerances. Where allowable "end-to-end" performance requirementsare not met due to interchanging components of the Apollo X G & N, the parametervariations will be used with a G & N error model and computer analyzed for missionperformance degradation.

Interchangeability Equipment Required

A complete Apollo X G & N System, including onboard spares and redundant units,will be required. The units of the G & N will be of the QA Class A Manufacturing status.

The support equipment required includes a complete GSE to equip a universalG & N, ISS-OSS, and test site, and one complete set of GFP, GSE, GOE auxiliaryequipment, and site equipment to support a universal test site. Integrated usage ofinfrequently used calibration equipment with acceptance and delivery testing may befeasible.

Interchangeability Testing Facilities Requirement

To support the Interchangeability testing a minimum of 20 x 20 feet of floor spaceof test area is required. Other requirements are identical to those previously listedfor the Acceptance and Delivery Testing Facilities.

QUALIFICATION TESTING

Qualification testing of the Apollo X G & N is based on the philosophy of "similarity"with the Lunar Apollo command module equipment. A comparison of the Apollo X andLunar Apollo mission operating environments and stresses reveals that they are basicallythe same except for flight duration. Therefore, the extent of testing has been limited to:

• Unique environments of Apollo X,

• Dynamic testing of significant structural or mounting variations from the LunarApollo equipment.

The entire philosophy of the qualification testing is to use the most severe environ-ment or stress where a choice between similar environments exists. Per NAA request,no provisions have been made for formal G & N reliability, life, or design safety factortesting.

-97- SID 64-1860-14

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SPARK PLUG - the Eladronlcs Division^ of GENERAL MOTORS CORPORATION

Qualification Objectives

The main goal of the qualification program for the Apollo X G & N is to assure ]that the environmental design considerations of the G & N have been met with respectto the mission requirements. The main objectives are as follows. ^

I• Demonstration that the unique structural configurations of Apollo X G & N will '

withstand the dynamic stresses of ND 1002037 without causing equipment degrada-tion. !

• Evaluation of the possible effects of material contamination or material inter-action with the two-gas atmosphere reflecting into adverse equipment perform- --jance. i

• Demonstration of the capabilities of Apollo X modules and the full G & N to per-form within specified tolerances when subjected to the temperature, temperature- 'altitude, and humidity exposures as detailed in ND 1002037 with the addition of a50-percent oxygen, 50-percent nitrogen atmosphere during the exposures.

Ii

At present, NASA utilizes ambient earth atmosphere as a gas environment. If theLunar Apollo qualification is eventually modified to reflect 100-percent oxygen, testing _^in the two-gas atmosphere exposure proposed herein could be extensively reduced. «

Qualification Testing Plan Implementation ~*i

Table 7 lists the applicable paragraphs of ND 1002037 that will be performed onApollo X G & N equipment. The Apollo X qualification program was derived with the ~]following assumptions.

• Results from the Lunar Apollo Program concerning electromagnetic interferenceand susceptibility will serve to qualify the Apollo X. s

• With the exceptions of the onboard spare IMU, PSA, and ECU packages, theApollo X G & N equipment will be qualified for mission vibration, acceleration, ~]and shock by the Lunar Apollo qualification program and normal manufacturingcontrols.

iThe entire Apollo X G & N qualification program will take place at AC Spark Plug.

The individual flowgrams, Figures 36 through 40, show the testing to be accomplished. ^The arrangement is only tentative, should facility usage cause conflicts, the order oftesting may be rearranged.

r—»Due to the length of time before qualification testing is to begin, it does not appear

feasible to detail the performance and test configurations of the PSA, ECU, D & C group, -and spare IMU. Rather, it is proposed to submit detail test planning and proceduresfor each unit prior to conducting the qualification test for that unit. A sufficient time jin the scheduling has been allowed for this activity.

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SPARK PlUC-th. Electronics Division of GENERAL MOTORS CORPORATION

General Equipment Performance

• Before and after environmental exposures, the G & N equipment will operateas specified by applicable FTM and JDC.

• During the environmental exposure those parameters and signals monitored willbe within the JDC tolerance except where requirements are specifically modifiedby engineering analysis to account for instrumentation limitations.

• No servo loop will exhibit any tendency to oscillate or lose control.

• During dynamic environments, all switches and relays will remain in their in-tended position and not produce inter mittents.

• The dynamic environmental exposures will not destroy or impair the ability ofthe G & N equipment to retain repeatable alignments, gimbal drift rates, gimbaltransient responses, and inertial component characteristics and coefficients.The degree of repeatability will be determined by engineering analysis.

• During all qualification testing, the G & N equipment will operate to keep thetemperature of inertial components between 120° and 140° F.

Specific Equipment Performance and Test Configurations

Module Performance. Before and after environmental exposures, the modules willoperate as specified by their applicable FTM's. However, during the exposure, signalsmonitored will be within the FTM or engineering-specified tolerances.

Module Test Configuration. Each module will be operating or nonoperating during theexposure,as specified. Upon completion of the environmental exposure, an abbreviatedfunctional test and visual inspection will be conducted before the module is removedfrom the test area for complete inspection and acceptance.

G & N System Performance. Before and after test exposures, the G & N will satisfactorilycomplete:

• Standby Power-On Test,

• Operate Power-On Test,

• Failure Indicating Circuitry Test,

• Temperature Control Test,

• G & N System Power Supplies Test,

• Guidance Computer Operational Test,

-10!-,- SID 64-18GO-14

Thlft document contain l&Jozinatlon*«£t0lotlntrtt«£natlonal def«na« of th« United SUtea wlthtath« m»inln« at tb* uploamfe l«w«. TuJ^Bg^S. C. Section* 793 ud 794. th« truumlulanor rmlttlan of which In uy m«fi5r to u^tuthorlud pereon u prohflUMl by Uw.

SPARK PLUG- th«Ele<tron!cs Division of GENERAL MOTORS CORPORATION

_106_ SID 64-1860-14

t• Optics Positional Accuracy Test (modified),

• Guidance Computer Mode Control, ; ,|

• Frequency and Step Response Test (modified),h• IRIG Scale Factor and Gimbal Torque Test, \ j

• PIP A Scale Factor and Bias Test,

• IRIG Coefficient Determination Test and Simulated Space Fine Align, ; j

• Gyrocompass ing Test.

During the exposure, as many tests as practical will be completed from the abovelist.

G & N Test Configuration. The Apollo X G & N will be installed in the applicable testchamber, with the optics aligned to a precision heading. Exercise of the G & N equip- —?ment will be accomplished by remote means to conduct the testing during the test ex- :, \posures. This will be accomplished by control with a remote control display keyboard,operation through the computer test set, and remote control switch actuators. •<-"*

i

Qualification Data Requirements^—^^^^B«^«~—_^^~MM~^B. I ' X

i i

Environmental Data

• Temperature— chamber walls * \i

• Pressure

• Coolant temperatures and flow rates : <

• Two-gas atmosphere composition — 50-percent oxygen and 50-percent nitrogen

• Humidity 1

• Random vibration inputs

• Acceleration inputs '

G & N Equipment Signal List "~"|i i

The signals listed in Table 8 do not necessarily constitute the normally availablesignals, but are only a preliminary estimate of qualification testing monitored signals flat varlouH times during the tout cyclo. ' '

Thll document contain! Information Biding thenpfronal detenie of the United States withinth« meaning of the espionage laws, TltlejCSi^TC. Sections 703 and TM, the tran»ml»lonor revelation of which In any manntc*=!S^nun*»Slu>rt£ed peraon u prohlhltad by law.

SPARK PLUG-the Electronics Division of GENERAL MOTORS CORPORATION

Table 8. G & N Equipment Signal List

Data List

+ 28 vdc Standby+28vdc Optics and JMU Operate+120 vdc+3 2 vdc+ 20 vdc-20 vdc-28 vdc2Hvac, 3200 cps4 vac, .1200 cps2vac, 3200 cps28 vac, 2-phase, 800 cps Wheel and Blower Voltages28v, 1-percent, 800 cps28 v, 5-pcrcent, 800 cps

PIPA Loop

X, Y, 7 Velocity Increment Pulses to Guidance ComputerX, Y, 7 Torquing CurrentsX, Y, 7 Precision Voltage ReferencesX, Y, 7 Scale Factor VoltagesPIPA Average TemperatureX, Y, 7 G/S OutputX, Y, 7 Preamplifier Output

ECU

4 vdcA() Pulses from (Juidance Computer (2 l inos)A0 Pulses to Guidance Computer (2 linos)Coarse Align Signals to Gimbal Servo A m p l i f i e rSignal to FDAIS1VH InterfaceGuidance Computer ECU EnableGuidance Computer Coarse Align EnableGuidance Computer Error Angle Counter Enable

Optics Loops

Motor Drive Amplif ier Outputs (4)Tachometer Outputs (4)X and Y Star Tracker Servo Outputs (2)Photometer Preamplifier Outputs (1)Photometer Output (1)Trunnion 2>: Resolver SignalsShaft and Trunnion Command SignalsStar Presence to ComputerHand Controller OutputsOptics Moding Switches OutputsComputing Hesolvors Outputs

Stabilization Loops

X, Y, 7 Gyro Preamplif ier OutputsX, Y, 7 Gyro Error Resolver OutputsIG, MG, OG Torque Motor CurrentsIG, MG, OG Coarse Al ign Signal f rom ECU'sIG, MG, OG \y and Kix Resolver Outputs to ECUX, Y, 7 A0 Pulses from Guidance ComputerCommon Gyro PVRX, Y, 7 Gyro Torquing CurrentGyro Average Temperature

-107- SID 64-1860-14

Thli document conUlu Information aifetho rooming of the ntplonage Uwa, Title 18,or revelation of which In any manne J""

fenee of the United Sutea withintlona 793 and 79<. the tranamlaalon

i«d person la prohibited by law.

SPARK PLUC thtJIertfOTlcs Dlvlito^ o| GENERAL MOTORS CORPORATION

Data Handling

Signals monitored will be recorded by one of the following: ! j

• Photoelectric recorders,•"1• Pen and Ink recorders, ' j

• Magnetic tape recorders,

• Paper printouts from data processing equipment,

• Logbooks and data sheets.n

The photoelectric, pen and ink, and written records will be reviewed, analyzed, \and summarized by Engineering. Magnetic tapes will be processed, and some signalsplayed back through G & N analog loop simulations to obtain gimbal and stable member Htransmissibilitles. All random vibration inputs to equipment will be processed through .' iappropriate filters and graphical presentations prepared.

n> »Qualification Equipment Required

• Test Specimens — Class A manufacturing status Apollo X configuration: '""I. \

• One Apollo X G & N system less onboard spares, including a Guidance Com-puter and two display keyboards, ->

• One each of every type module in the PSA (not including ECU modules), '

• One Apollo X spare IMU, j-\t

• One Apollo X PSA with all modules installed, '

• One Apollo X redundant ECU package, —j

• One Apollo X D & C Group. •' '

• Support Equipment H

• One Full Universal Test station GSE complement " ^

• Instrumentation Requirements H

• Thermocouple temperature recorders,

• Chamber pressure gages, ^

• Wet and dry bulb temperature recorders,

-108- SID 64-1860-14

n

Thl« document contains informationa&»<,^i£ the jta^Kal defense of the United State* withinthe iM..»t»g of the espionage laws. Title 18j^^0^^Sectlan« 793 and 794, the tranamlBilonor revelation of which In any tnann^^^an unaa&&ci««d person 1« prohibited by law.

SPARK PLUG-the Electronics Division of GENERAL MOTORS CORPORATION

• Mass spectrometer,

• Magnetic tape data acquisition equipment,

• Random noise analyzer,

• Signal conditioning circuitry,

• X-Y plotting equipment,

• CEC recorders and amplifiers,

• Sanborn recorders and amplifiers.

• Special Fixturing and Cabling:

• Environmental GSE-PSA simulated junction box,

• Remote switch actuators,

• Vibration and shock holding fixtures (4),

• Cabling to mate with temperature-altitude

f eedthrough connectors and test specimens and GSE.

• Facility Requirements:

• Temperature-vacuum chamber —NRC space simulator,

• Temperature-vacuum chamber with simulated star and solar sources,

• Explosion chamber,

• Vibration exciter,

• Shock exciter,

• Approximately 50 x 50 feet of test area,

• Gas handling facilities.

-109- SID 64-1860-14

This document contain* lnform*tion*SB«$^ng thjMlfonal defence of the United States withinthe meaning of the evplonage lava, TitleIO££^C.Sectlanj 793 and 794. the tranamlMlonor revelation of which In any manne^loan ur3te»riied person l» prohibited by law.

SPARK PLUG-the Electronic* Division of GENERAL MO CORPORATION n

p,

nn-

SID 64-1860-14

This document contain* infornutloaCh« n^f«ninj gf ({,0 •gpioQAge Uwf, Tltlojj

or nrelitlon of which In «ny

dofen.. of the United SUU* wtthlltcnm 793 ud 704. the trmumluloaIted per»oo li prohibited bv l«w.

SPARK PLUG-tht Electronics Division of GENERAL MOTORS CORPORATION

V. DEVELOPMENT PLANNING

OBJECTIVE

This section presents a brief and comprehensive description of the proposedApollo X Program. This effort includes design determination, design support, manu-facture, test, and field support of the Apollo X G & N hardware as described and im-plied in the technical portion of this study.

A major objective of the work effort will be to utilize as much as possible, with-out impairment of mission objectives and noninterferences with the Lunar Apollo Pro-gram, the design already completed for use in the Lunar Apollo Program. The majorstress on design changes will be to increase and sustain reliability of intended Apollo Xmissions,

SCOPE

AC Spark Plug will determine the Apollo X hardware design, and will completethe development of the equipment to be defined during the Program Definition Phaseof the project. Continued support will be given to the items of the Apollo X Programcommon to the Lunar Apollo Program.

The contractor will procure, fabricate, and deliver hardware to be defined bynegotiated work statements. Required testing during the various stages of manufactur-ing will be accomplished.

Due to the developmental nature and the preliminary planning stage of the Apollo XProgram, the schedules and general tasks presented constitute projections and are con-sidered to be beyond the control of AC Spark Plug. Specific control and definition isexpected to be established via a firm proposal during the Program Definition Phase.

The above activities will be supported by such groups as Management, Reliabilityand Quality Control, Engineering, Documentation, and Field Operations. The generalscope of activities for these groups follows.

-Ill- SID G4-1S()0-14

Thle document contain* Inforfliattqj^ffecting the^fatlonal defence of the (Jolted State* withinthe meaning of the espionage iawa. ntftHfegpTc. Sections 793 and 794. the tranenUailanor revelation of which In any iMrffs^toiut'tiaajpthortzed percon la prohibited bv l»w

SPARK PlUG-the Electronic! Division of GENERAL MOTORS CORPORATION

MANAGEMENT—)

The Management function will include such items as program technical control, 'issuance of proper directives, utilization of PERT and hardware delivery schedulecontrols, handling general program financial control, and maintaining liaison with '' \NASA, NAA, and other agencies and subcontractors. '

AC Spark Plug will provide and control documentation as specified in anticipated ""!negotiated work statements as a portion of the overall management effort.

RELIABILITY AND QUALITY CONTROL . I

Reliability and Quality Control have as a basic responsibility the continuation of pireliability and quality efforts already established during the Lunar Apollo Program. iThese are in general conformance with NASA Reliability Publication NPC 250-1(July 1963) and MIT/IL Reliability Plan R-349 (Revision B). The quality control pro- 1gram will be in general conformance with NASA Quality Assurance Publication __ .NPC 200-2 (April 1962) and MIT/IL Quality Assurance Plan R-396 (Revision A).

i—•»

An additional responsibility is to adapt the above techniques and guideline docu- 'ments to unique areas of the Apollo X Program.

. !

ENGINEERING '

~^Engineering will undertake as its portion of the overall program function such ;

tasks as follow.

• Sufficient and necessary design analysis of the Lunar Apollo equipment to enable "~^a firm and detail determination of common design usage with Apollo X configura-tion.

• Complete design analysis and verification of performance of the Apollo X equip-ment.

• Detail and sustaining design effort comprising engineering and technical support ~required to enable the selection of engineering models (assemblies, parts, andcomponents) that will meet the requirements of the design analysis and lendthemselves to production. This effort includes engineering support to productionand packaging design. •

• Complete performance, procurement, assembly and test, and operating proce-dureB for the Apollo X G &. N equipment. <

• Performance of syaLern tetits and checkouts, development test support, develop-ment of necessary test setups, and assembly and test of the complete Apollo X ?G & N equipment.

-112- SID (54-1860-14

Thli document contalni InfonnatlonateH&ngtjjfraitlonal defonae of the United SUtea withinthe "M^mtng of the eaplonage lawa.^gJft^STSt^S. C. Sections 793 and 794. the tranamlaatonor revelation of which In any manner to anunxuthorlEed peraon la prohibited bv law

SPARK PLUG-tin Electronic Division MOTORS CORPORATION

FIELD OPERATIONS

In addition to the tasks specified above, AC Spark Plug will conduct a Field Opera-tions Program for the Apollo X G & N equipment. This program will provide super-vision, administration, and planning for field site activities, support activities, spareparts, associated test equipment, as well as site operation, personnel for the operationand support of the contractor-furnished equipment from factory acceptance throughpost-flight analysis.

TEST SCHEDULES AND MILESTONES

To accomplish the general objectives outlined on an integrated but noninterferencebasis with the Lunar Apollo Program, the Apollo X major milestones and the Apollo Xintegrated equipment and test schedules, presented in Figures 41 and 42, weredeveloped.

These schedules and milestones are set forth as an integral effort on uniqueApollo X hardware configurations and basic Apollo configurations. Alteration of deliveryconfigurations would required re-appraisal of the milestones and the schedules.

-11:5- SID G4-18GO-14

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gectlng thg^iatlonal defenla of the United State* withinj(Tlc. Section. 793 and 794. the tranamlMlon

orlied person li prohibited by la».

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Rlonal detente of the United State* withinC-JSectiona 70S and 794. the tranamlMlon

orlzed peraon 10 prohibited bv law

SPARK PlUG-the Elactron JDivlslim of JENERAL MOTORS CORPORATION

SID 64-1860-14

Thi» document contain. lnformatnS%lM*g the national defenee at the United 8ute» withinth. meaning of the enplonage l»w«. Tltlel^^^C. Section. 793 and 794, the truumUilonor revelailoB of which In any manner to| a&^Bkuthorlzed perion u prohibited by law

SPARK PlUG-the Electronics Division of GENERAL MOTORS CORPORATION

VI. CONCLUSIONS AND RECOMMENDATIONS

CONCLUSIONS

The Apollo Block II G & N System can meet all the LPO mission performance re-quirements with the addition of the recommended modifications, spares, and redundancy.

The LPO measurement uncertainty in Altitude after one orbit is well within theorbital altitude accuracy requirements of 1.1 nm using either ground tracking, staroccultation, or landmark tracking.

The G & N System can provide a reference to maintain the spacecraft attitude towithin 0.125 degree of local vertical and the spacecraft rate to within 0. 02 degree/sec. of the desired rate about all axes.

The G & N System, with the recommended modifications can meet an overall relia-bility of 0.9975 for the LPO mission with 91 pounds of spares and redundancy. A i^elia-bility of 0. 9965 can be attained with 65 pounds of spares only. Further improvementscan be expected after completion of those areas where further engineering developmentis necessary.

RECOMMENDATIONS

When Block II design is completed, perform a complete parts stress analysisbased on circuit application of parts.

Implement life testing of Block II mechanical and electromechanical componentsunder conditions similar to their Apollo X applications to determine life and to obtainrealistic stress factor ratings as recommended in the test plans.

Update the Block II failure effects analysis, presently under study, to the Apollo Xconfiguration to determine which part and module failures can cause a G & N Systemfailure in the mission.

Define alternate G & N capabilities for performing required functions after thefailure effects analysis has shown that a module failure can produce an adverse G & Neffect.

Investigate the effects on Apollo X system reliability of repeated on-off powerswitching .

-117- SID 64-1860-14

This document contains InformSttea^flectlnaJpe national defence of the United States withinthe meaning of the espionage laws. TlT^^^S. C. Sections 793 and 794, the transmissionor revelation of which In any manB^tote-ujauthoriied person Is prohibited by law.

SPARK PLUG-tha Electronics Dlviilw of GENERAL MOTORS CORPORATION

n

SID 64-1860-14

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or r*v«ttllan of which In any n

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iorl»~l perion It proh|bll*d hy Uw.

'•AC SPARK PlUC-tht El»t«ronl(f Divisional GENERAL MOTORS CORPORATIONI.X ••»}

VII. APPENDIX

SUPPORTING DATA

Table A. 1 is a detailed Apollo X G& N Time Line for the lunar polar orbit mappingmission, earth polar orbit mapping mission, and the low-inclination earth orbit missionphases. This table is a further breakdown of Table 1 of this report.

Table A. 2 lists the component and subassembly reliability failure rates for theApollo Block II G Si N. These values were used to compute the probability of successvalues given in the Technical Analysis portion of this report. An asterisk (*) is usedto denote the main contributors to the failure rate, and a dagger (t) is used to denotethe secondary contributors.

There are a limited number of built-in alternates for performing the requiredG & N functions as shown in Figures A. 1 through A. 11. In many instances, the samemodules are used, but in alternate modes. A detailed failure effects analysis wouldindicate whether an alternate approach is available after a module failure has occurredand what the resultant improvement in reliability would be.

-119- SID 64-1860-14

C. Scotlou 793 ux) 7M. Uw trM.mi.llooTfcli document conttla* InformiSItk» •—*-«-f af tb« Mptouce l«w«or rvnUtlon of which la my mlnner

SPARK PlUC-lftt Ekctnmlrs Olvliion of GENERA I MOTORS CORPORATION

Table A. 1. Apollo X —G & N Time Line

Time (minutes) Event

0-12

12-2020-25

2S-175175-180

180-192192-197

197-227227-257

257-815615-8.10

830-847847

647-3215.1215-32:10

3230-32473247

3247-43404.140-4.155

4.155-44424442-4447

4447-44574457

4457-45004500-45054505-45174517-4521

4521-4641

4841-4848

4646-48581856-4657

4657-4781

4761-4788

4766-47704776-4777

4777-4870

4870-4875

4875-48814881-4941

4941-5228

5226-5241

5241-52565256-52H6

5286-5571S57I-5SHO

r.5sr,-r,fioir.«o i-5n6i566I-M61M-ifi-.vioi.1961-5976f.97r,-6(|06

6291-

Launch Into 100 nm Parking OrhltGuidance Monitor

Alignment Update

Alignment Update

Translunar Inject Guidance Monitor

Dock CSM to lab

Alignment

Mid-Course Correction

Alignment

Mid-Course Correction

Alignment

Alignment Update

Mid-Course Correction

Alignment Update

Deboost to liinar Orbit

10-Star Occultatfon Measurements

Alignment Update

Orbit Correction

9-Star Occultallon Measuroment

Alignment Update

Orbit Correction

8-Star Occulintlnn Measurement

Alignment Update

I-ocal Vertical M;ip

Alignment

l.ocal Vortical Map

Alignment

Iflcal Vertlciil Map

Alignment

l/ical Vertical Map

Mission PhaseInertlal

SubsystemOptical

SubsystemOutdanceComputer

PROFILE I

Boost S-Vi4T

Earth Orbit

1

fTrnnslunar Inject

1

Translunar

\

First Lunar Orbit

Second Orbit

Thin) Orbit

One-Half Orhlt

One -Fourth Orbit

One-Half Orbit

One-Fourth Orbit

•

IIT

,iii

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

41 • |r'f' Itepeat alxwe mapping cycle for a total of 5r.-half orbit maps and 55-quarter orhlt maps except a fu l laliimmcnt w i l l precede the f irst local vertic.il map of each cycle.

-120- SID 64-1860-14

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of which In uiy m«n^ri«TO5l»«uUiorl««l p>rioti u proklbllxl b)r l.w.

SPARK PLUG-lh« EUttronfci DMilw tjI GENERAL MOTORS CORPORATION

Table A. 1. Apollo X — G & N Time Line (cont)

Time (minutes) Event Mission PhaseIncrtlnl

SubsystemOptical

SubsystemGuidanceComputer

PROFILE I (cont)

44. 166-44,46644.451-44,466

44,466-44,48144,481-44,484

44.484-44,52644,526-44,531

44,531-44,54144, .14 1-44, 543

44,543-45,51045,510-45,525

45,525-45,54045, 540

45,540-47, 11047,110-47,125

47,125-47,14047, 14047, 140-48,670

48,690-48,70548,705-48,720

48,720-48,72548,725-48,740

48,74048,740-48,785

48,785-48,79048,790-48,804

48,804-48,822

Alignment

Injection

Alignment Update

Trajectory Plane Change

Alignment

Mid-Course Correction

Alignment

Mid-Course Correction

Alignment

Alignment Update

Mid-Course Correction

Alignment Update

Entry

Transearth Inject

1

Trim

1

Enlr

searth

1

y

I

I1

1

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

PROFILE II

0-10

10-20

20-27

27-55

55-60

60-66

66-68

68-128

128-143

143-158158-218

218-503503-518

518-533533-563

563-848848-863

863-878878-T

T

(T<14) -<T«35)

I _Monitor launch to 100mm Parking Orbit Boost S-V and SIVT)

S1VB Relight (392 seconds)

Transpose, Dock and Const

Alignment Update

T. V. C.

Orbit Verification and Corrections

Alignment-Update

Local Vertical

Alignment

Local Vertical

Alignment

Recycle 158 to 87H as often as nil. nil

T. V. C. (18 seconds)

EntryT = 720C t 158; C = Total Cycles

Parking Orbit

Transfer Trajector

Earth Orbit Uijoctlo

First Orbit

Second Orbit

ows

Debixist

Entry

'

n

1i

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

Tkl> doounmtIk* owulaf Ol tk* <»yloa>c* Uwi.or rmUtloo <rf which In any

-121-

tlon.I d«I«ai. <rf UM United SUtM wlUil«8. c. 8««lon» TM ud TM. I!M Inuumlulea

Ki Mrion If Dnklbltod br Uw.

SID 64-1860-14

SPARK PlUG-tht El««ronl« Diviilon of GENERAL MOTORS CORPORATION

Table A. 1. Apollo X— G & N Time Line (concluded)

Tlmo (minutes) Event Mlmlon PhaseInertia!

SuhHyslemOptical

SubsystemGuidanceComputer

PROFILE III

0-10

10-20

20-50CA git>MJ"nO

85-90n/\ t fintfU — I UU

100

100-125

125-130

i irt i AnUV~ 1 'til

140

140-200

9nrt i i T7rt£VU~.I I * if ('J

31,370-31,430

11 Jin ii 7ftn.1 I , T.IU""') 1 » ( I f U

3 1,790-31, 80S

31,805-31,82031.820

31, 820-31, 880

3 1 880-64 490

64,490-64,550

A4 fi*yi A4 flirtO4» O.TU""1!! » iU

64,910-64,925

64 92r>-64 94064,940fid ttdft Ad QIAn^t a'tlt— O^ t UO^

64,954-64,975

Monitor Launch Into 80 mm Pirklng Orbit Booit S1VD

Dock C8M to I.ob

Alignment Update

SM Light for 10 sccondn

Alignment Update

8M Light for 10 secande

Orbit Verification and Corrections

O ft N Checkout

Alignment

SM Light

Orbit Verification

O & N Checkout

Alignment

SM Light for 18 seconds

Entry

Orbit

Orbit Transfers

200 mm Orbit

Orbit Correction

Dobooat

Entry

T

1X

IX

J

1

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

—!

.1; i

-122- SID 64-1860-14

Thl« docom«« rooUiaj laformfttea.jUectliuy^Miitionil ttltat* al UM Uolt«d State* within<*• mm>»« o(tk* uplonmg* Uwi. - 115 1 :8. c. 8,01 or,. iu mi 7M. th« tnaimlulinor rwrtlUlOB of wtlch In uiy oMfiSr to^^antuthorln..! por.on I. prdilblud by IM.

$PABK PLUClhtIectronlcf Olvl.lon oUENtRAl MOTORS CORPORATION

Table A. 2. Block II Module Failure Rates

Assembly

INERTIAL MEASUREMENT UNIT

Electrical Connector

Slip Ring*

Temperature Alarm

Temperature Control

Ducosyn Transformer

Thermostat

Padding Resistors

Heaters

Hesolver*

Bearing*I

Torque Motor*

16 PIPA Preamplifier*

16 PIP*

PIP Suspension*

IHIO and Preamplifier*

IRIG Suspension*

Precision Resolve r Adjustment

Blowers and Controls

Water Connector

POWER AND SERVO ASSEMBLY

PIPA Calibration*

Gyro dc Calibration

Pulse Torque Power Supply

AC Differential Amplifier and Integrator

Gfmhal Servo and Alignment Amplifier*

Fall Indicator and Warning*

3200 cps AAC Filter*

3200 cps, 1-percent Amplifier*

-28 vdc Power Supply!

Sextant Motor Drive Amplifier

Scanning Telescope Motor Drive Amplifier

"00 cps Optics Comparator

Servo Integrate Relay

Antl- Creep

Scanning Telescope Trunnion Transformerand Relay

Failures perin' Hours

0.02

3.0

0. 1

0.25

0. 2f,

0.06

0. 0!>

0. 1

5.0

0. (!

5.0

1.55

4. 2

2.0

12.0

2.0

1.4-1

2.4

0. 2

2. 2

6.6

8.8

7.3

6.0

7.3

2.8

8.5

3.9

2.2

2.2

0.8

12.0

3.fi

6. 2

linage

16

6

1

1

1

4

3

4

4

6

3

3

3

3

3

3

1

2

2

3

1

1

3

3

I

1

1

1

2

2

1

I

1

1

Assembly

POWER AND SERVO ASSEMBLY (cunt)

Trucker Modo Relay

Sextant Mode Relay

DC Differential Amplifier and PrecisionVoltage Reference*

Binary Current Switch*

O fc N Subsystem Filter

ROOcps AAC Filter*

SOOcps, 1-percent Amplifier*

SOOcps, 5-pcrcont Amplifier*

ELECTRONIC COUPLING UNIT

Quadrant Selector*

Cos (9 -i/') Generator*

Coarse System Electronics*

Main Sum Amplifier and Quadrature*

Digltol-to-Analog Converter*

nimlral IflRlc'

Error Angle Count and Logic

Clock and Mode Logic*

4-Volt Power Supplyt

DISPLAYS AND CONTROIS

Control Assembly

G 6 N Indicator Control Panel

Attitude Impulse Switch

Sextant Hand Controller

Assembly Parts

aPTICS HEAD

Resolvor*

Motor Generator*

F.lectrlcal Connector

Bearing

Transformer

Resistor

Failures per10* Hour*

12.0

fi.O

2.7

2.6

0 .2

2.7

3.5

3.7

17.7

17. 1

8.2

11.8

4.5

5.9

6.0

5.6

1.7

4. 26

6.0

12.4

13. 24

5.0

5.0

0.2

0.6

0.2

0. 1

Usago

1

1

4

4

3

2

2

3 .

5

5

5

5

5

5

5

1

1

1

7

1

1

5

4

10

51

1

2

-123- SID 64-1860-14

This document contain<k* BMStaf of th* Mptonac* Uor rmlttlOB of which In

aBeqfJaTthe national d«fana» of UM United Mate* vithlai. U.8.C. Sectlou 793 utd m. UM tnnsmlutoa

m«nn«r»-^ji uaauthorluil txr»» i. r,,v*n,i._< w i—

> SPARK PLUG-flit Electronics Divhlon of GENERAL MOTORS CORPORATION

"73 ^3 g3 "CQ ••*

T

a>66.

<<oj-,a

"]

1

n

j~*

SID 64-18(50-14

or r#»«lttlOD of which in any

nation*] del«w« of UM UolUd BUU>* within. C. Seel ion a 703 ud 194. th« lrmn»ml*«iao

[•uthorlrvd parion !• prohibited toy lw».

m SPAIK PlUG-lhi Electronics Division of GENERAL MOTORS CORPORATION

r ~>1

Attitude 1 fc Zero 1Control ECU w "

Coarse AlignToS/C

Zero Optics _ _ _ SXT Follows ^. _AGC ft* AGC "°-*

U ' 11 1

Zero _ _J IOptics | 1 h— —

1

11 .

i *i[_^ Manual — ^

Direct

SXT FollowsTracker

ManualKesolved

A '

VV ~~| * ~~* Align ""*"

—I "1

i i1

• ManualDirect

ManualResolved i

L i

Figure A. 2. IMU Alignment — Update

1 1

Attitude 1 Zero 1Control * *" ECU

ZeroOptics • • fc SXT Follows _ -AGC | [ AGC

i 1

SXT FollowsTracker

ij

Zero _J (

Optics ^ """Manual

Resolved

L*

|_ _ ManualDirect ~T"^

ManualDirect

ManualResolved

0 ^ Fine

i Align

— ~1 "

W

_TiiiiUx

*y 1

1. . - __ ,

1 1

Figure A. 3. IMU Alignment — Update

_125_ SID 64-1860-14

Thl§ doeumot eoctUtM Infonnatloo^fcitttlot th;.a*»«i5il deftnn c( U» UmlUd gutn wttklmIk* m«ii)lTH <* Uw nploiuc« lawi, TIU^^^TC. 8*ctl<»« 7§S od T94. th« tnumlMloi

SPARK PLUG-the Electronics Division of GENERAL MOTORS CORPORATION

AttitudeControl

Manual S/COrientation

T1

Figure A. 4. Thrust Vector Control

Main S/COrientation

Figure A. 5. Midcourse Correction —Orbit Correction

Zero

Optics iMnnual1)1 r<;ct.

1

~!

LJ

Figure A. 6. Landmark Tracking

Figure A. 7. Attitude Control

AttitudeControl

Figure A. 8. Boost Monitor

1Figure A. 9. Nothing

-126- S1D (»4-18()0-H

ThJa document contain* informatlotht mMLolnc of the e*plontgeor revelatlOD of which in any

) defen«a of the United 6Ut*» withinSection* 793 ud 794. the tran«ml*alonhortrrd oerntm ll prnhtbltx) hv law

SPARK PLUG-the Electronic! Division of GENERAL MOTORS CORPORATION

Figure A-10. Star-Landmark Measurement

Manual S/COrientation

AttitudeControl

Zero OpticsAGC

ZeroOptics

SXT Follows1 t ' AGC ? **

1 I

1 1 1

1 1 11 1 11 1 11 | *•*•

1 1

!i

iiSXT Follows

Tracker

ManualResolved

ManualDirect

1T

\t/AI

1I|1

i

r{<Ji

t

•NJ

i

\ f

* ManualResolved

1*

11

Tracker

Figure A. 11. Star-Landmark MeasuremcMit\

-127- SID (>4-1860-14

Tola document contain* Informthe meaning of the eeplonage laws,or revelation of which In any

/national defense of the United States within0". S.C. Sectiona 793 and 794, the tranamlulon