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  • 8/8/2019 Apollo Experience Report Guidance and Control Systems Mission Control Programer for Unmanned Apollo Missions


    C A L N O T EI


    4, - A N D APOLLO 6F. Hollozuay

    B. Johnson Spdce CeuterTexas 77058

    A E R O N A U T I C S A N D S P A C E A D M I N I S T R A T I O N W A S H I N G T O N , 0. C . J U L Y 1 9 7 5

  • 8/8/2019 Apollo Experience Report Guidance and Control Systems Mission Control Programer for Unmanned Apollo Missions



    2. Government Accession No.

    7. Author(s1;ene F. Holloway

    17. Key Words (Suggested by Author(s))Unmanned Mission .Control Systems.Command Guidance.Spacecraft Guidance

    .Automatic Flight Control

    9. Performing Organization Name and Address

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

    ,yndon B. Johnson Spacecraft Center[ouston, Texas 77058

    19. Security Classif. (of this report)unciassiiiea

    20. Security Classif. (of this page)TT-.-.l n -";#;,.,A 6021. NO . of Pages_ _ " ic. IU " "*I A bU

    2. Sponsoring Agency Name and Addressrational Aeronautics and Space Administrationiashington, D. C. 20546

    22. Price'$4.25

    5. Supplementary Notes

    3. Recipient's Catalog No.5. Report DateJulv 1975

    ~ 6. Performing Organization CodeJSC- 04214

    8. Performing Organization Report No.S- 43210. Work Unit No953- 36-00- 00- 7211. Contract or Grant No.

    13. Type of Report and Period CoveredTechnical' Note14. Sponsoring Agency Code

    6. AbmactLn unmanned test flight prog ram w as requir ed to evaluate the command module heat shield andhe st ru ct ur al integrit y of the command and serv ice module/Saturn launch vehicle. The missionontrol programer was developed to provide the unmanned interface between the guidance and navi-;ation computer and the other spacecra ft s yste ms fo r mission event sequencing and real- time groundontrol during miss ions AS-202, Apollo 4, and Apollo 6. The development of th is unmanned pro-;ramer is tra ced fr om the initial concept through the flight tes t phase. Detailed discussions of hard-rare development problems a r e given with the resulting solutions. The mission control prog ra me runctioned corr ec tly without any flight anomal ies fo r al l miss ions. The Apollo 4 mission cont rolirogramer was reused for the Apollo 6 flight, thus being one of the fi rs t subsystems to be reflownIn an Apollo sp ace flight.

  • 8/8/2019 Apollo Experience Report Guidance and Control Systems Mission Control Programer for Unmanned Apollo Missions



    Section PageSUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1MISSION CONTROL PROGRAMER DESCRIPTION . . . . . . . . . . . . . . . . . 2

    3nput Keying Commands fo r Mission Sequencing . . . . . . . . . . . . . . . . .8

    15192 1

    DESIGN USING EXISTING TECHNOLOGY . . . . . . . . . . . . . . . . . . . . . 2 3Spacecraft Command Controller . . . . . . . . . . . . . . . . . . . . . . . . . 23Ground Command Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . 28Attitude and Decelera tion Senso r . . . . . . . . . . . . . . . . . . . . . . . . . 29

    Sequencing To Accomplish Mission Requirements . . . . . . . . . . . . . . . .Real-Time Commands for Ground Control . . . . . . . . . . . . . . . . . . . .Backup 0 . 0 5 g Acceleration Sensor . . . . . . . . . . . . . . . . . . . . . . . .Sequencing Postlanding Recovery Aids . . . . . . . . . . . . . . . . . . . . . .

    TEST EQUIPMENT DESIGN . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30DEVELOPMENT SCHEDULES AND TEST PROGRAM . . . . . . . . . . . . . . . 32

    Brea dboard and Proto type Development . . . . . . . . . . . . . . . . . . . . . 333 3

    Te st Equipment Certification . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 4Production Delivery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35Qualification Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35Interface Verification Te sts . . . . . . . . . . . . . . . . . . . . . . . . . . . 39

    Electrom agnetic Int erf erence Considerations . . . . . . . . . . . . . . . . . .

    Spacecraft Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 1HARDWARE PROBLEMS AND RESOLUTIONS . . . . . . . . . . . . . . . . . . . 43

    Reliability an d Quality O bjectives . . . . . . . . . . . . . . . . . . . . . . . . 43


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    Section PageRelay Failures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43Solder Contamination in MCP Relays . . . . . . . . . . . . . . . . . . . . . . . 44Polarized Tantalum Capacitor Failures . . . . . . . . . . . . . . . . . . . . . . 45Time-Delay Failures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45Cracking of Glass Seals Caused by Clipping of Relay Pi ns . . . . . . . . . . . . 49

    FLIGHT PERFORMANCE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50Mission AS-202 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50Apollo 4 Mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51Apollo 6 Mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51

    52Unmanned Flight Planning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52Development Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52Test and Test Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52Hardware Problem s . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53

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



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    Table PageI APOLLO 4 MISSION DISCRETE EVENTS SUMMARY . . . . . . . . . . . 10

    I1 NOMINAL MISSION SEQUENCE OF EVENTS FOR SECONDSPSFIRING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13III REAL-TIMECOMMANDS . . . . . . . . . . . . . . . . . . . . . . . . . . 16IV NOMINAL. MISSION SEQUENCE OF EVENTS FROM ENTRY TOLANDING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 0V NOMINAL MISSION RECOVERY SEQUENCE OF EVENTS . . . . . . . . . 2 1

    VI RELAY FAILURE HISTORY . . . . . . . . . . . . . . . . . . . . . . . . . 44FIGURES

    Figure1 Mission control programer . . . . . . . . . . . . . . . . . . . . . . . . . .2 Spacecraft command controller . . . . . . . . . . . . . . . . . . . . . .3 Attitude and deceleration sensor . . . . . . . . . . . . . . . . . . . . . .4 Block diagram of the MCP . . . . . . . . . . . . . . . . . . . . . . . . .5 Diagram of the G & N sys tem connector interface with the SCC . . . . . .6 Diagram of the S-IVB IU int erf ace with the SCC . . . . . . . . . . . . . .7 Di agr am of launch control and GSE inte rface with the SCC . . . . . . . .



    8 Circuit logic and switching relays(a) Simplex (not redundant) . . . . . . . . . . . . . . . . . . . . . . . . . 23(b ) Dual series (redundant) . . . . . . . . . . . . . . . . . . . . . . . . . 2 3(c) Dual parallel (redundant) . . . . . . . . . . . . . . . . . . . . . . . . 2 3(d ) Dual series. triply parallel (redundant) . . . . . . . . . . . . . . . . 2 3

    9 Cable assembly wire harn ess . . . . . . . . . . . . . . . . . . . . . . . . 2510 Printed wiring board with components . . . . . . . . . . . . . . . . . . . 26


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    Figure Page11 Bracket showing re la ys and time-delay mountings . . . . . . . . . . . . . 2712 Automatic checkout equipment inter face req uir ement s fo r theMCP redundancy test . . . . . . . . . . . . . . . . . . . . . . . . . . . 2813 Qualification testi ng sequence . . . . . . . . . . . . . . . . . . . . . . . 3614 Fractured solder terminals

    (a) Example1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38(b) Example2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38(c) Example3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3815 Repair method. soldering backside of ter mi nal str ip

    (a) Orientation of rework are a . . . . . . . . . . . . . . . . . . . . . . .(b) Closeup of rework area . . . . . . . . . . . . . . . . . . . . . . . . .


    16 Typical diode mounting bracke t . . . . . . . . . . . . . . . . . . . . . . . 4017 Time-delay circuit schemati c . . . . . . . . . . . . . . . . . . . . . . . . 4618 Relay schematics to desc ribe glass-seal problem

    (a) Bottom view of relay-case header . . . . . . . . . . . . . . . . . . . 50(b) Top view of rel ay p ins with cont acts in open configuration(c) Cr os s section of pin entering relay header and the gla ss seal . . . . .. . . . . 5050


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    FOR UN MA N N ED M I S S O N S A S -202, A POLLO 4, AND APOLLO 6By Gene F. Hol lowayL y n d o n B . J o h n s o n S pa ce C e n t e r

    S U M M A R YThe unmanned Apollo missions AS-202, Apollo 4, and Apollo 6 were successfulflights.missions, were accomplished fo r each mission. The mission control pro gra me r unitwas successfully used for all three missions without causing a flight anomaly o r thelo ss of any functional event fo r which the pro gra mer was respons ible. The missioncontrol progr ame r met all the flight and ground test objectives without the loss o rerr one ous indication of any nec essa ry output. The miss ion control pr og ra me r did,however, expe rience individual component failures during the pro gra m. These fewfail ures wer e compensated for in the redundant circuit design of the mission controlpr og ra me r and did not resu lt in the los s of or deficiency in any necessary missionoutput.

    and a cr ew was not pres ent to compensate for possible flight anoma lies by switching toalternate backup systems o r by using alte rnat e mission modes, the mission controlprogramer with its som et ime s trip ly redundant paths was r equi red to have higher inher-ent reliability than other Apollu sys tem s. The Apollo 4 mission control progr ame r w a sreflown during the Apollo 6 mission.reflown on a space-flight mission.

    The flight objectives, which were a prerequi site f or the manned Apollo

    Because the mission control pro gra me r was designed fo r unmanned missions

    This unit was the first Apollo sy ste m to be

    INTRODUCTIONThe st ru ct ur e and heat-shield design of the Apollo command and serv ic e module

    (CSM) had to be verifi ed under Saturn V launch and lunar - ree ntr y e nvironment s beforeit could be considered man- rated. The missio n control pr og ra me r (MCP) was devel-oped by the NASA and the CSM pr im e contractor to prov ide the automatic event switch-ing interface between the input command and control systems (e. g., the guidance andnavigation (G&N) computer ) and the output response s ys te ms f o r the Apollo unmannedtest flights.atti tude cont rol and sequencing, The objective of thi s repor t is to document the MCPdevelopment program from the initial concepts and mission requirements phase; throughthe des ign and manufactu ring buildup testing; during the spac ecraft installat ion and

    The MCP al so provided the real-time ground-control inte rfac e fo r backup

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    te st s; and, finally, through the launch, recovery, and postflight anal ysis . This Apolloexperience resulted in useful information that should be adapted to the design of fu tur eunmanned space-flight equipment.The unmanned flight req uiremen ts for the MCP were identified by the interfac ingsubsystem design engineer s and ground flight controllers. This rep ort gives a missiontime line for the Apollo 4 mission and demonstra tes how this mission was accomplishedusing a few key commands from the G&N computer and using the int erna l logic andhardwired time delays of the MCP to drive o r switch the interfacing spacec raft s yst ems.The backup ground-control capability is listed together with a description of each real-ti me ground command. An example is given to show how the ground commands could beused to provide a backup thrust maneuver.The re quirement to te st each redundant path o r sys tem in the Apollo launch vehiclejust before launch was an essen tial fact or in the mission s uc ce sse s of the Apollo Pro-gra m. This report discuss es the problems that had to be resolved to perf orm thes espacec raft redundancy t es ts on the MCP.During the MCP development, changes to the sp acecraf t were approved thatrequ ired design changes to the MCP. Some of the spa cecraft changes are listed in thisrepor t together with thei r effect on the MCP design.A s an aid to the rea der, where neces sary the original units of me asu re have beenconverted to the equivalent value in the SystGme Inte rnat ional d'Unit6s (SI). The SI unitsare written first, and the original units are written parenthetically thereafter.


    The MCP (fig. 1)consist ed of three units: the spa ce cra ft command cont roll erThese units wer e located i n the space-(SCC; f ig . 2 ) , the ground command control ler (GCC; si mi la r to the SCC), and theattitude and deceleration se ns or (ADS; fig. 3).cra ft on a platfo rm assembly mounted in plac e of the cre w couches on the cr ew couchshock mounts.(360 pounds) so that it could provide the weight ne cessary to verify the resp onse of thecre w couch st ru ts to landing impact s. A s shown in figure 4, the keying commandswere supplied to the MCP by the G & N sys tem , the Satu rn IVB (S-IVB) instrumentationunit (IU), the updata link, and the launch cont rol complex.sys tem s actuated by the MCP output switching functions ar e a ls o shown in figu re 4.

    The MCP weight was adjusted to approximately 163 kilograms

    The other interfacing


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    Mai n display console

    Spacecraft command _ - - -

    Attitude and decelerationGround command

    Auxiliary power systemci rcu i t breaker boxCrew compartment

    access hatc h outli ne 'Components of missi oncontrol programer(typical)

    Figure 1. - Mission control prog ram er .Input Keying Commands for Mission Sequencing

    The pri ma ry sou rc e of m ission sequencing key commands to the MCP was theG&N sys tem computer.were as follows. The original keying commands furnished by the G&N computer

    1 . G&N abort2. Positive- or negative- Z antenna switching3 . Flight di rec tor attitude indicator alinement4. Gimbal motors5. G&N fail6. 0.05g7. Positive-X translation


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    8. Command module (CM) and service module (SM) separation9. G & N entry mode

    10. G & N change in velocity A V mode11. G & N attitude control mode

    Note: Spacecraft command con tro ller shown:ground command controller similar

    Figure 2. - Spacecraft command controlle r.The first two interface signals, G & N abor t and positive- o r negative- Z antenna switch-ing, we re removed fr om the G & N wiring on spacecraft 017 and 020 because the failuresthat could produce e ithe r signal wer e considered t o be single-point fai lure s. Thedecision was made that, because the abort signal o r relay closure could be erroneous,the G&N system computer should not automatically abo rt a mission. Because real- t imeground commands were ava ilable to swit ch the antennas and because the G & N systemcomputer controlled the s pacecr aft attitude during the missio n midc ourse flightpath,


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    the sof tware tas k of pro graming the G & N system computer to switch antennas automat-ically w as considered too costly fo r the resu lts achieved. The diag ram of the connectorinterface between the MCP and the G & N system computer (fig. 5) shows that the MCPprovided the 28-V dc power supply for the G & N system relays, and the G & N systemcomputer provided the logic and relay clos ures to complete the circui t paths.

    -Attitude switch

    Attitude switch


    Pushbutton switch


    Figure 3 . - Attitude and deceleration senso r.The S - I V B I U provided fou r keying commands fo r the MCP. The following listrepresents the S-IVB inte rface keying commands fo r mission sequencing. Each

    command was dually redundant.1. S - I V B res ta r t A2. S-IVB es ta r t B3 . Launch escape tower jettison command A4. Launch escape tower jettison command B5. Lift-off signal A


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    6. Lift-off signal B7. Launch vehicle and spa cecraft separatio n start A8. Launch vehicle and spacec raft separation st ar t B

    Whenever the spacecraf t direc t- cu rre nt bus w a s powered, the MCP provided redundantdirect-current power to the S-IVB I U fo r the generation of the d is cr et e sequencingsignals (fig. 6) .internal time-delay circ uits to provide the other required m ission sequences.These keying sign als fr om the IU were used by the MCP ogic and

    Thrus t vector cont rolAtt i tude control

    Thrust vector andReaction controlcont rol system system system indic ation

    Figure 4. Block diagram of the MCP.

    A diagram of the logic circu itr y of the int erf ace between the MCP and the launchcontrol and ground support equipment (GSE) is shown in figure 7. This interface pro-vided launch- control personnel with the capability to di sa rm the pyrotechn ics, switchoff the logic buses, and ope rate the onboard flight rec or de rs while the spacecra ft andlaunch vehicle were stacked at the launch site. The prog ram re se t signal of this con-trol interface allowed launch-control pers onnel to start the MCP; hat is, to reset alllatching relays and pre par e the MCP ogic circuit for lift-off.


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    It 2 8 V d c a r m e dpr ior to launchI


    t2 8 V dc armed byCM-SM separate

    +28 V dc armedby launch vehicle/spacecraft separ atecommand t2.5 seci


    NCcI -C













    + oG&NFlight director attitude indicator alinet2 8 V dc;imbal moto rs ON v128 V dc %- r - -G&N failure indication ON v+28 V dcSpare0.059 indication-". -+X translat ion

    CM-SM separate ON

    N CG&N entry mode+28 V dcG&N AV mode+2 8 V dc1

    &N attitude control mode128 V dc

    MCP connectoraON except after launch escape system abort or G&N failbThe C-28, etcet era, nomenclatu re ref ers to relays associated withthe Apollo guidance computer.



    bC-2 8





    C 24


    D14K40K 14K3 9

    D14K38D12K4101 2K40



    Figure 5. - Diagram of the G&N system connector interface with the SCC.


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    time references tg and tC and between t4 and t6 listed in table I and described intable 11. These sequences, which were in itiated by the G& N system computer, a r e notlisted as specific times (hours, minutes, and seconds) but are given as referencesymbols (tA, tg, t2, etc.). The detailed software prog ram s for the G&N-systemcomputer established these specific times fo r the various missions.

    (2 )

    ( 3 )

    + 2 8 V dc bus A

    +28 V dcSpare launch

    ' contro l p ins

    Launch controlinput to MCP







    No connection













    L- To ground support equipmentl launch control

    Command power A 1Master events sequence controllerlMESCl pyro bus A safe commandMESC pyro bus B safe command

    MESC pyro buses A&B arm com

    MESC logic bus A safe command

    MESC logic bus 6 safe command

    MESC logic buses A&B arm command

    Tape recorders record

    Tape recorders stop recorder

    Tape recorders rewind

    Start operation of environmental control systemCalibrate telemetry

    Program reset

    Command control transfer

    Master control transfer

    Pyro =pyrotechnic


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    Saturn V ascent to orbitGuidance reference releaseLift-offSaturn IC (S-IC) inboard engineS-IC outboard engine cut-offSaturn I1 (S-II) engine ignitionS-IC in ter sta ge jettisonLaunch e scape syst em jettisonS-I1 engine cut-offS-IVB engine ignitionbS-IVB engine cut-off



    Earth parking orbitStart Earth parking orb itStart second orbit revolution

    Second S-IVB firing bS-IVB engine ignitionbS-IVB engine cut-off(SPS) firingBegin reorienta tion to cold-soakattitudeEnd of reorienta tion to cold-soakatti udeCSM/S-IVB separation

    Coast to fi rs t ser vice propulsion syst em

    ~~ ~ ~~

    Planned tim efr om lift-off,h r :min: se c(4

    00: 00: 16 .7000: 00: 00.0000: 02: 15.5000: 02: 32.4 000: 02: 35.2000: 03: 04.3500: 03: 08.3500: 08: 39.5500: 08: 44.0500: 11 :05 .40

    00: 11:15.6001: 38: 20.00

    0 3 : 11:54.5003: 17: 12.53

    03: 17: 27.71

    03: 20: 42.8103: 27: 14.43

    Actual time,h r :min: sec

    - -00: 00 : 00.26300: 02: 15.5200: 02: 30.77

    ------00: 08: 39.7600: 08: 40.7200: 11:0 5 . 6 4

    00: 11: 15 .6--

    03: 11:26.603: 16:26.3

    03: 26: 28.2aThe planned times given are taken fro m AS-501 Spacecr aft Opera-tional Traj ecto ry, Volume I- rajectory Description, August 25, 1967.bRefers to guidance signal.

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    Begin reorientation to fir st SPSEnd of reorientation to fi rs t SPS

    ignition attitudeignition attitude

    Fi rs t SPS firing bSPS engine ignitionbSPS engine cut-off

    Earth intersecting coastBegin reorientation to cold-soakEnd of reorientation to cold-soakApogeeBegin reorientation to second SPSEnd of reorien tat ion to second SPSReaction control system (RCS)


    ignition attitudeignition attitudethrusters on

    Second SPS firing bSecond SPS engine ignitionbSecond SPS ,engine cut-off

    Planned timefr om lift-off,h r :min: sec(4

    03: 27: 22.7303: 27: 51.81

    03: 28 : 52.7303: 29: 18.93

    03: 29: 24.6803:29: 53 .7605:49:04.3208: 01: 36.7508: 02: 01.0508: 14: 40.42

    08: 15: 10.42

    08: 19: 34.40

    Actual time,hr :min: sec

    03: 28: 06.603: 28: 22.6

    --= tB (note c )= tC (note c)

    08: 10 : 54.8= t 4 (note c )08: 15: 35.4= t6 (note c)

    The planned ti mes given a r e taken fro m AS-501 Spacecraft Opera-tional Trajectory, Volume I- rajecto ry Description, August 25, 1967.a

    bRefers to guidance signal.Table I1 provides additional information.


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    Event~~~ ~

    Preentry sequenceBegin reorienta tion to CM/SMEnd of attitude orientation, coas tCM/SM separationStart CM attitude orientation fo rEnd of attitude orienta tion, coas t

    separa tion attitudeto CM/SM separation

    entryto entry

    Atmospheric entry0.05g indication121 9 2 0 - m ( 4 0 0 000 f t ) altitudeEnter S-band blackoutEnter C-band blackoutExit C-band blackoutExit S-band blackoutEnter S-band blackoutExi t S-band blackoutDrogue-parachute deploymen.Main-parachute deploymentC M landing

    Planned timefro m lift-off,h r :min: se c(a)

    08: 20: 12.9708: 20: 54.0108: 2 2: 07.8508: 22: 12.8508: 22: 36.03

    - -08: 23: 35.0208: 23: 57.0008: 24: 01.0008: 25: 55.0008: 26: 19.0008: 30: 15.0008 : 31: 47.0008: 35: 39.0008: 36: 27.0008: 41: 25.00

    Actual time ,hr :min: se c

    08: 18: 02.6--

    08: 18: 06.28

    08: 19: 56.28----

    --08: 31: 18.6--- -

    %he planned tim es given a r e taken fro m A S - 5 0 1 Spacecraft Opera-;ional Trajectory, Volume I- rajectory Description, August 25, 1967.

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    Time Initiatedreference by- outputto-unctionCPfunction

    tg + 1.0 sec


    ' T/C~ spsI sps

    Reorientation to second SPS ignition attitudeG &NG & NG & NG &NG & N





    Monitor mode OFF and G&N attitudeInitiate pitch maneuverComplete pitch maneuverFlight director attitude indicatorFlight director attitude indicator

    control mode ON

    aline ONaline OFF

    Second SPS thr ust maneuverX



    G&N attitude control mode OFF andMonitor mode OFF and G& N A VPositive-X translation ON

    monitor mode ONmode ON

    Gimbal motors ONEntry batteries to main dc busesFlight qualification recorder O NPrepilot valve A ONData storage equipment rec ord er ONPrepilot valve B ONY a w 1 gimbal motor start




    a'Electrical power syst em.

    Stabilization and.contro1 sys tem.c;Transmitting and control function.


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    Time Initiated M C Preference 1 by- 1 function 1 Function outputI to-I I I ISecond SPS th rust maneuver - Concluded

    t 3 + 1 . 5 sect 3 + 2 .0 sect t- 2 .5 sec3t 3 + 3.0 sect 4t 5t6

    t6 + 3.0 sec






    Y a w 1 gimbal motor ONPitch 1 gimbal motor st ar tPitch 1 gimbal motor ONY a w 2 gimbal motor startY aw 2 gimbal motor ONPitch 2 gimbal motor startPitch 2 gimbal motor ONSPS thrus t O N (second firing)Positive-X translation OFFSPS thrust OFFGimbal motors OF FGimbal motors OF FRemove entry batteries f rom mainSelect third gimbal position se tPrepilot valve A O FFPrepilot valve B OF FG & N AV mode OFF and monitor


    mode ON


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    The MCP w a s designed with the specification that the initiation ti me s fo r parti c-ular keying and sequencing commands for performing various mission functions couldbe changed from mission to mission. However, the detailed integrated sequence ofevents to accomplish any particular mission function would remain consistent for allmission s. Table I1 lists the functions required to reo rient the space craft to the secondSPS engine ignition attitude and the functions required to initiate and complete thesecond thrust maneuver. The M C P time delays ar e shown in the time-re ferencecolumn. For example, "t + 1.0 sec " indicates that the "yaw 1 gimbal motor start"3signal fro m the MCP to the SPS gimbal actuator motor oc cur red 1. 0 second aft er theG&N system computer had given the "gimbal motors ON " signal (t3) to the M C P . Thetime-delay un its were hardwired, potted plug- in modules that were hermet ical ly sealedin a metal case. Several time-delay selections were available for a given module baseconnection size. Because of the high-start - cur rent requ irement s of the motors, thegimbal motors were turned on at 0. 5-second intervals to prevent an ele ctri cal overload.Table I1 gives a function- by-function descrip tion of the spa cecraf t system activityrequire d to per form a n SPS thrust maneuver. A sim il ar functional listing can beobtained for all the required missi on events, including abor t sequencing.

    A t time tg, the MCP turned on the flight qualification r ec or de r and the data-storage - equipment reco rder . The sequences of events that were considered the mostcritical or of the highest pr ior ity and that were to be recorded varied significantlyfr om mission to mission. The reco rder s and cameras had a limited tape and filmcapacity ; and timed on- off sequences, which varied significantly fr om miss ion tomission , wer e nec ess ary to obtain only the most important data. These changes insequence ti me s required MCP hardware changes. Usually, the times var ied s o muchthat different connector interface ci rcu its had to be selected fo r the keying commands(e. g., a command for a second SPS fi ring gimbal motors off ins tead of a command fo rS-IVB/spacecraft separation). These MCP hardware modifications were costly interms of money and schedule time, requiring new engineering drawings, specifi cationrevisions, te st equipment modifications, recert ification of the te st equipment, andreacceptance tes ting of the flight hardware. In fu ture manned o r unmanned develop-mental flight prog rams, strong emphasi s should be given to the developmental instru-mentation inter face with the spacecraft systems. For launch vehicles or spacecraftthat contain flight computers, the instrumentation used to monitor flight events duringthe developmental program should be designed so that the changes of instrumentationsequences from missi on to mission can be placed in the er asab le portion of the comput-er memory. If th is procedu re were followed, the sequences could be quickly and cheap-ly modified in r eal time.

    R e a l - T i m e C o m m a n d s for G r o u n d C o n t r o lThe MCP, through the GCC unit, provided the switching logic ci rcu it ry , therelays, the relay driv ers, and the required spacecraft syst em interface and had thecapability to proc ess 77 ground- commanded signals received by the spac ecra ft throughthe digital updata link. This technique provided a backup perf ormance capability to thespacecraft by using ground support personnel and their flight control consoles to providethe updata-link signal commands. A list of the tit les and number codes of po ssib le real-time cnmmands is given i ~ t a h l p11; The number codes corr espond to the appropri ate flightcontrol console switches in the Mission Control Center for the uplinked transmissions.


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    Real- t imecommandnumber

    0100070 60 20 30 40510111 2131 415161 7202 122232 42526273233

    %ot used.

    1 6


    A b o r t light (system A ) ONaAbort light (system A) OFFaAbort light (system B) ONaAbort light ( system B) OFFaFuel cell 1purgeFuel cell 2 purgeFuel cell 3 purgeReset real-time command numbers 0 2 to 0 4Lifting entryDirect thrust ONDirect thrust OF FReset real-time command numbers 10 to 1 2Positive pitch dir ect rotationNegative pitch direct rotationPositive y aw direct rotationNegative yaw direct rotationPositive rol l direct rotationNegative roll d ire ct rotationDirect ullageReset real-time commands 1 4 to 22Propellant OF F SM quad APropellant OFF SM quad BPropellant OF F S M quad CPropellant OF F SM quad DPropellant O N SM quad APropellant ON SM quad B


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    TABLE 111. - REAL-TIME COMMANDS - ContinuedReal- timecommandnumber


    TitlePropellant O N SM quad CPropell ant ON SM quad DLaunch escape tower jettisonG&N f a i lG& N f a i l inhibitReset real-time command numbers 41 to 42Roll rate backupPitch r at e backupY aw rate backupFlight direc tor attitude indicator alineReset real-time commands 44 o 47Negative- Z antenna ON (very- high-frequency (vhf)Positive- Z antenna ON (vhf sc im it ar only)Roll A and C channel disableR o l l B and D channel disablePitch channel disableYaw channel disableReset real-time command numbers 54 o 57CM and SM separationUpdata link S-band receiver selectUpdata link ultrahigh- f r equency receiverHydrogen tank 2 heater fan ONOxygen tank 2 heater fan ONHydrogen tank 1 heater fan ONOxygen tank 1 heater fan ONReset real-time command numbers 64 o 67

    scimitar only)



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    TABLE 111.- REAL-TIME COMMANDS - Concluded~~



    Launch escape tower ab ort and MCP se parationSpareC-band OFFC-band ON (2 pulse )vhf transmitter OFFvhf transmitter ON

    Real-time commands 14 to 21, 23, and 54 to 60 were to be used to control thespacecr aft attitude if the au tomati c attitude control provided by the G&N sy st em had notfunctioned properly. If th is malfunction had occur red, the automati c channels to thereaction control system (RCS) could have been disconnected by real- time commands54 to 60, and the di re ct rotation commands, rea l-t ime commands 14 to 21, could havebeen transmitted. The dire ct rotation commands required that the ground con trol lerstran smi t the time interval neces sary to achieve the des ire d spacecraft attitude. Forexample, if the G&N sys tem had failed, the second SPS fir ing sequence could have beenaccomplished by ground controllers using real- tim e commands according to th e follow-ing sequence.1. Send real-time command 41, "G&N fail.2. Use real-time commands 14 to 2 1 to position the vehicle to the pro pe r firin gattitude.3. Send real-time command 11, "direct thrust ON, I t at the desired firing time toautomatically st ar t the gimbal mot ors in sequence and to initiate the firing.4. Monitor the spacecraft trajectory by using the Mission Control Center real-time tracking data.5. Send real-time command 12 , "direct thru st OFF, I t at some predeterminedveloci ty point o r at the violation of a limit line on the tra ject ory plot.6. The vehicle probably would be orie nted fo r CM and SM separa tion by thereal-tim e commands cited in step 2.example that the G & N syste m had failed; theref ore, additional mis sion objectives wouldnot be attempted.

    Such a probab ility would be consistent with the


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    7. Send real-time command 61, "CM and SM separation, I ' to ar m the masterevents sequence- controll er (MESC) logic circuitry, to a r m the pyrotechnic devices,and to initiate the separa tion sequence.8. The vehicle would be oriented fo r entry.The advantages of re taining a degr ee of real-t ime ground cont rol of the spac ecra ftwould have been demonstrated if a spacecraft system failure had actually occurredduring one of the flights. If the G&N syst em h ad malfunctioned, som e useful heat-shielddata at the requ ired high-entry velocit ies sti ll could have been obtained. Seve ral com-binations of s pac ec ra ft sequencing and control, other than that of the G&N syste m mal-function, could be accomplished by real- tim e commands. The flight operation plansand the launch r ule s for e ach missio n furnish a descr iption of the many possibl e alter-nate missio n modes. Severa l real- time- command numbers are intentionally omittedfrom table 111.the Apollo 4 and 6 mission s.The following real-time commands were deleted by the NASA before

    Number Title30 CM RCS system A propellant OFF31 CM RCS system B propellant OFF36 CM RCS sy st em A propell ant ON37 CM RCS system B propellant ON53 G&N antenna switching

    B a c k u p 0.05g A c c e l e ra t io n S e n s o rSevera l significant mission events were required between the e ntry phase and thelanding (table IV).0.05g deceleration is reached, which occ urre d at an altitude of approximately88 400 me te rs (290 000 feet) is a cri tica l mission event fo r reco very of th e spac ecraf t.The pr im ary determina tion of the ent ry point (0.05g) was made by the G&N sys tem,and a redundant 0.05g s ignal was provided in case the G&N sys tem failed to providethis signal or in c as e the G&N sys tem had failed earlier in the mission.signal was produced by ac cel erom eter s in the ADS unit of the MCP. Table IV showsthe MCP tr ans fe rr ing the 0.05g signal fro m the G&N syst em to the stabilization and

    control syste m (SCS) at t The importance of accurately determining the point atwhich 0.05g was reached cannot be overemphasized, because, after th is point is passed,the method of controlling the spac ecra ft with the RCS th ru st er s is changed. The pitc hand yaw attitude control was inhibited by the SCS, and the spa cecraf t was st eer ed byusing the RCS th ru st er s to roll the space craft about a n offset cen ter of gravity. Otherimportant spacecraft systems (e. g. , the Earth-landing sys tem (ELS)) were al so acti-vated when the 0.05g point was sensed. Thus, t he MCP per fo rmed an importa nt func-tion on unmanned fligh ts as the redundant deceleration indicator.

    The se nsi ng of the point of a tmosphe ric en tr y (the point at which

    This redundant



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    Timereference Outputby- function to-unctionnitiated MCP

    to (about 28 m inbefore landing)

    G &N Xii 1ESC

    7620-m (25 000 ft ) altitude


    t (maximum of0 13 min beforelanding)

    to + 20 sec

    t + 270 sec0

    EL S


    0.05g ON = 88 400 m (z 90 000 ft)0.05g signal (backed up by the MCPEarth landing s yste m (ELS) activate AELS activate B7620-m (25 000 ft) barometric switch armedSwitch to negative- 2 antenna

    0.05g backup function)





    7620-m (25 000 ft) baro metr ic switch activatedSCS/RCS enable OFF


    I C Mpex cover jettisonedDrogue-parachute deployment (r eef ed)Drogue-parachutes disreefedA r m 3658-m (12 000 f t ) barometric switch


    3658-m (12 000 f t ) altitude3658-m (12 000 ft) barometric switches A and B

    Sta rt landing backup 14-min ti me rConnect C battery to flight and postlandingRCS fuel dump activate ARCS fuel dump activate BArm landing switchvhf recovery beacon O Nvhf survival beacon ONRCS purge activate ARCS purge activate BImpact landing

    (F&PL) bus




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    S e q u e n c i n g P o s t l a n d i n g R e c o v e r y Ai dsAnother function of the MCP w a s to sequence the postlanding recovery aids for theunmanned missions. The cor rec t performance of thes e functions was necessary toensu re the recovery of the spacecraf t after landing was successfully achieved. Thesequence of even ts afte r landing (table V) w a s initiated by the impact of the space craf t

    on the water. The impact was sensed by triply redundant switch acc elerome ters in theADS unit of th e MCP. The MCP was als o requ ired to te st and certif y the uprightingsyst em of th e spa cecr aft before a manned flight . The ADS unit contained triply redun-dant attitude indicators that could sense whether the spacecraft w a s floating apex up(stable I) or apex down (stable 11). If a stable I1 signal had been indicated by the attitudeswitches (table V ), logic ci rcuit s in the SCC would have re laye d a signal to the upright-ing sys tem to inflate the flotation bags.

    TABLE V . - NOMINAL MISSION RECOVERY SEQUENCE OF EVENTS[Al l events are MCP funct ions init iated by the MCP.]


    Function OutputI to-to

    t + 11 sec0

    Nominal events~~

    Impact landingMain-parachute disconnect AMain-parachute disconnect BArm attitude indicatorConnect entry ba tteries to F&PL busConnect auxiliary batter ies 1 and 3Connect auxiliary batteri es 2 and 3Remove entry batteries fr om main busesDeploy high-frequency (hf) recoveryMESC logic bus A safeMESC logic bus B safeFlashing light O Nhf tra nsc eiv er ON (stable I only)Circuit breaker 45 OPEN

    to F&PL busto F&PL bus

    antenna (stable I only)



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  • 8/8/2019 Apollo Experience Report Guidance and Control Systems Mission Control Programer for Unmanned Apollo Missions


    D E S I G N U S I N G EX IST ING TECHNOLOGYThe critical development schedules for the MCP required that existing electronicComponents that had been previously qualified on other missile o rechnology be used.space prog ram s were selecte d whenever practicable.

    Spacecra f t Com mand Con t ro l l e rThe MCP block diagram in figure 4shows the Apollo system interfacesre quired by the MCP. The SCC unit of theMCP provided the logic capability neededto accomplish the interfac e and event-sequencing requ irements . The event-sequencing and switching functions for the

    unmanned flights were accomplished bythe use of relays . Thes e hermeticallysealed microminiature general-purposerelays, which had an all- welded construc-tion, were used extensively in the logicand switching circ uitry. The rela ysoperated at 28 V dc and had a 2-, 3-, or10-ampere cur rent rating.The redundancy requi rements of theMCP were classifi ed into four categories.1. Simplex (not redundant)- heMCP output o r rea l- time- command func-tion may fail either ON o r OFF because ofa single MCP component failure(fig. 8(a)).2. Dual se ri es (redundant)- TheMCP output or real - time- command func-tion shall not fail ON as a result of anysing le MCP component fa ilure (fig. 8(b)).3 . Dual parallel (redundant)- heMCP output or real-time-commdnd func-tion shall not fail OFF as a re su lt of anysingle MCP component fa ilu re (fig. 8(c)).4. Dual series, triply parallel(redundant)- he MCP output or real-time- command function mus t respond cor-rec tly in the event of a single MCPrnmpnnent fai111re (fig, El!d)),

    Inputa V d c't output(a) Simplex (not redundant).


    (b) Dual ser ie s (redundant).

    o u t p u t T . . I 28VdcI Relay A Input1 35 1

    ( c ) Dual parallel (redundant).2 8 V dcInput

    Relay A

    1 6'Re la y B

    (d ) Dual series, triply parallel (redundant).Figure 8. - Circ uit logic and switching rel ays .


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    The redundancy requirements for the MCP design were established by using thesefour categories. A request for a definition of the requi rements of the sys tem int er facewas submitted to the appr opri ate engineering design groups, and spec ific redundancyrequire men ts were obtained on a n event- by-event basis fo r numerous potential mis-sions.redundancy requirements.The design of the MCP was then established consistent with these missi on

    The redundancy options that were used in the MCP are shown in figure 8. Exam-ples of equivalent redundancy could als o be i llu st ra ted within the MCP, howing the us eof time delays, capaci tors , diodes, et cetera.because they repr ese nt the majority of the components in the MCP.The relays are used for illustration

    The cir cuit s in figure 8 are shown with rel ay contacts configured in the no rmallyopen state .normall y closed state.design.as long as the switching signal is applied to the solenoid.the switched configuration until a n additional res et switching signal is applied to thereset solenoid of the rel ay.

    Similar redundant configurations are used with the relay contacts in aBoth momentary and latching relays were used in the MCPMomentary r elays remain switched into the changed- sta te configuration onlyLatching relays remain in

    A trip ly redundant grounding network was used throughout the cable- ha rn es s andpanel-harness assemb lies (fig, 9) to provide elec tric al grounds for the MCP.grounding scheme was important in accomplishing the bench tests and spacecraft teststhat verified the redundant components within the MCP. During tests, thes e grounds(Gl, G2, and G ) wer e alte rnat ely cycled (opened and closed) or cycled in combinations(G G , G2G3, o r G1G2) to isol ate and veri fy the operat ion of specif ic redundant paths .For example, in figu re 8(d), as su me relay A ope ra tes with ground G * relay B withground G ; and relay C with ground G3. The redundant paths would be verified asfollows.


    31 3


    1. Make contact between grounds GI and G and leave G open.2. With the prope r signal to the relay solenoids, re lay s A and B close their

    27 3contacts.

    3. Step 2 verifies the c ente r path shown in figure 8(d). The top and bottom cir-cuits remain open because relay C has no ground to complete i ts circ uit and does notclose.4. Make contact between grounds G2 and G3, and leave GI open.5. Relays B and C close the ir contacts, and rel ay A remains open.6. Step 5 verifies the bottom path shown on figur e 8(d). The two top paths remainopen because relay A has no electr ica l ground to activate its solenoid ci rcui ts and thecontacts do not close.7 . Relay A in the middle path does not have its contacts failed in a closed positionin st ep 3, because in step 6 the middle path opened. If, in step 6, the middle row of con-ta ct s had not opened, the fa il ur e of relay A in a closed position would have been indicated.


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    Figure 9.- Cable assembly wire harness.

    This method can be continued until each redundant function is verified. Thisredundancy was an important pa rt of the Apollo Program, because the proper function-ing of each redundant circuit path had to be verified just before the spacec raf t w aslaunched. The pro cedure of checking redundant ci rcui try before launch w a s used forthe Mercury spacecra ft and was continued fo r the Apollo spacecraft. In general, thisverification of redundant circ uit paths w a s a simple tas k fo r the Apollo spacecraft,because most sy st em s wer e designed to be dually redundant (system A and system B).The power could be removed fro m eit her system to verify the proper functioning of thecompanion sys tem. However, checking the redundant circuitry in the MCP became adifficult and tedious job because of the many complex ser ies-pa ra lle l cir cui t paths.As shown in figu re 2, the SCC used 1 9 printed wiring boards (control assemblies).

    The detailed logic cir cui ts, relays, time-delay circuits , and other components wereplugged into these control assembl ies ; the components of this c ircuitr y wer e standard-ized and interchangeable. For example, a 3-ampere latching relay could be inter -changed with a 3-ampere momentary relay, or a 15-second time-delay device could beinterchanged with a 60-second time-delay device (figs. 10 and 11). Great ca re had tobe taken by the manufacturing personnel when removing a componept that had previouslybeen mounted and sold ered to the printed wiring board; otherwise, the meta llic tra ckcould be lifted fr om th e board or damaged. The control assemblies used f or the vari-ous MCP systems were also standardized and interchangeable. For example, controlassembly 6 in MCP system 2 and the s imi lar control assembly inMCP syst em 4 couldbe interchanged to resolve a problem with solder closeout relays.The SCC had 1 5 connectors to meet t h e various interface requirements and toprovide sufficient test points f o r ground tes ts. The unit had a ground-shorting connec-tor and three GSE connectors that were instrumented fo r the box-level bench tests.The internal grounds could be automatically applied and removed while the operation ofvarious components was being verified on the bench test console.


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    Figure 10. - Printed wiring board with components.

    The importance of des ign flexibility must be emphasized. A description has beenHowever, thisgiven of the increased flexibility obtained by in terfacing the MCP with the G&N systemfor its input keying commands r athe r than by using fixed pr es et t imer s.increased flexibility w a s limited. The inte rface connectors of the G&N system, theS-IVB IU , and the launch control and GSE (figs. 5 to 7 and 12) provided the MCP withcapabilities fo r 1 5 different flight keying and sequencing commands, which could bemodified fo r each mission, and 1 2 prelaunch keying commands. For approximately120 different miss ion events, the MCP furni shed the logic cir cu itr y and internal timedelays f o r switching the output to the interfacing sys tems at the corre ct mission times.The capabilities of the hardw ired logic circ uit ry w e r e not as flexible as had been desired.As mentioned previously , changes in the miss ion event sequence on-off time s of inter-

    facing hardware so metim es resulted in major MCP design changes.The following ar e examples of MCP design changes resulting f rom changes i nmission plans o r in interfacing system requirements.1. The planned tra jec tor ies for the Apollo 4 mission indicated a possibility ofspacecraft skipout during the entry phase , The MCP originally had latching re lays toprevent the loss of the 0.05g signal once it w a s obtained. During the Apollo 4 mission,the 0.05g signal could be obtained, los t during skipout, and then obtained again; the re-fore, the latching relay had to be replaced with a momentary r elay .


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    Figure 11.- Bracket showing relays and time-display mountings.2. A 14-minute time delay was requi red aft er a sensing function (indicating anequivalent barometric pressure at a 3658-meter (12 000 foot) altitude) was added to the

    MCP. This function was an ELS backup to initiate cutting of the parachute sh rouds 5 to10 minutes after landing.3. The gimbal motor on-off time s were changed fro m mission to mission to pre -vent the actuator clutches from overheating.4. A time delay was added in the MCP o prevent damage to the high-frequencyantenna by not allowing the antenna to deploy before the spacecraft w a s in an apex-upattitude in the water.5. Before the mission, the SPS engine gimbal positions were pred icted f or eachfiring during the mission. These positions were pre se t in the MCP o prevent larg egimbal position changes and la rge transi ent s during the firi ng initialization. Each space-craft had different center- of- gravity requirements at the various firing times; thus,each spacecra ft required different gimbal position settings that necessitated modificationof the MCP system.


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    Ne w interface ,C 6 A l J 1 3 : I -Grou nd command G-1control e r dc negati veGro und command G-2 0

    -LI IE acontroller dc negative II J IIGrou nd command G-3controller dc negative

    controller dc negativecontroller dc negativeSpacecraft command c-3 zcontroller dc negative A-


    Figure 12. - Automatic checkout equipmentinterface requirements for the MCPredundancy test.

    6. On-off sequ ence s fo r the tape re-corde rs and came ras were changed for eachmission.The se few examples of th e hardware changesmade t o the MCP indicate the flexibilityrequi red of a developmental flight systemsuch as the MCP. For example, before thedes ign of the MCP was complete, prep lan-ning should have dete rmined that the gimbalposition settings would va ry fr om mission tomiss ion and that the har dware should bedesigned so that a technician could changethe se tt ings without opening the SCC. When-eve r this unit was opened, a complete reac-ceptance test was required. In futureprograms, the changeable characteristicsof unmanned developmental flight testsshould be recognized, and various flexiblesoftware methods of prog raming miss ionchanges should be considered.

    G r o u n d C o m m a n d C o n t r o l l e rThe GCC unit of the MCP interfaced primarily with the updata link. This unitprovided the switching-logic circuitr y, the relays, the relay driv ers , and othe r com-ponents for processing the 77 real- time ground-command signals originating at the flightcontrol consoles in the Mission Control Center.degree of flexibility re qui red by the SCC. During the program, GCC changes were made

    t o cor rec t design problems and to eliminate cer tain capabilities, rath er than to r eviseand redesign logic and interface circuitry . As previously discusse d, five real- tim ecommands were elimina ted from the Apollo 4 and 6 missions ; the GCC wiring associatedwith these commands was cut and stowed.

    The GCC design did not re qui re t he

    The GCC used component and wiring redundancy si mi la r to that previously de-scr ibed for the SCC. The series- redundant cir cui try (fig. 8) was the most commonlyused circuit logic; however, para llel - redundant circ uit ry was used for proce ssing thereset real-time command, and ser ies- para llel - redundant c ircu itry was used f or proc-essin g the abort command.The GCC was designed to respond to minimum cu rr en t inputs of 18 to 24 milli-

    am per es with a pulse duration of 25 to 3 5 millisec onds from the updata link. Also, theunit was designed not to respond to current levels less than o r equal to 28 milliampereswhen pulse durat ions were less than or equal to 1 millisecond. Ea rly electromagneticinterference (emi) tests at the factory showed that the GCC relay dri ver s wer e trigger-ing on noise voltages, and resistor-capacitor filter networks had to be added to eachrelay driver. This design change was the most significant fac tor incorporated in theGCC. The ge ne ra l configurat ion of the GCC and the SCC is the same (fig. 2).


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    The reset real-time commands 05, 13, 23, 43, 50, 60, and 70 were necessa rybecause, once the GCC relay driver received a minimum-value curr ent pulse from theupdata link, the associate d latching relays were activated. The real- time commandcould be removed o r canceled only by sending a rese t command. Some rea l- tim e com-mands used momentary re lay s (e. g., positive- Z antenna ON). These momenta ryrelays were on as long as the command was being transmitted and off at all other times.The use of latching re lay s saved the ele ctr ica l power that would have been req uired tohold the relay solenoid in the activated state and amounted to considerable power savingsfo r events that would be on for long periods.

    A t t i t u d e a n d D e c e l e r a t i o n S e n s o rThe ADS unit of the MCP performed the critical spacecraft r ecovery requir ementsduring the entry, landing, and reco very phase s of the mission. The ADS des ign (fig. 3)w a s simple, consist ing of the following maj or components.1. Three spring-mass impact switches (accelerometers) to sense the water

    impact during landing2. Three pendulum-mass attitude indicators to sen se stable I o r stable I1 orienta-tion of the spacecraf t after landing3. Two line ar a ccel ero mete rs to se nse the 0.05g level during entr y4. A pivot shaft and pivot fr am e for ground te st of the attitude indica tor s5. Push- to- test switches fo r ground testing the 0.05g and landing acc elero met ers6. A radio-frequency interference filter f o r the input powerComponents of the ADS were used in the following or de r during a mission.1. The 0.05g accel erom eter s were armed by a signal from the SCC at the t ime ofCM and SM separation. These accele rometer s were designed to tri gger at decelerationsof 0. l g to 0. 5g, a higher decelera tion value than the 0.05g value furnished by the G&Nsyst em. In Jun e 1966, the 0.4g spr ead in the tolerance of the backup dece lera tion sen-sor was recognized as possibly causing a wide deviation between the actu al and plannedspacecraft landing points when the backup signal w a s used. A specific te st wa s thenadded to the box-level acceptance test to measure and r ecor d the exact decelera tionleve l of th is sensor. For spacecra ft 017, thi s deceleration value w as 0.29g * 0.04g for

    initiating the backup 0.05g signal. The 0.05g signal would be automatically overr iddenby the ADS in c ase of a skipout trajecto ry. The *O. 04g tolerance could not be discardedbecause the ac cele rom eter s were te mperatur e sensitive and the precis e flight tempera-tures were not defined.magnitude c los er than the initial values.The trigger point was better defined, however, at an or der of2. The th ree impact switches were armed by a signal relayed f ro m the SCC whenit sensed an altitude of 3658 me te rs (12 000 feet). The impact switches were designednot t o trig ger fo r impact pulses le ss than 4.7g and to trigge r for values of approximately5g and above. A push-to-test switch was provided for each impact switch for ground-test purposes.


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    3. The thre e attitude se nso rs were a rm ed by the impact decele ration pulse,These s ens ors indicated stable I whenever the apex of the sp acecraft w a s approximatelyk65" from an upright position. When the apex of the spacecraf t dropped below the6 5 " point, stable I1 w a s indicated by the sensors . Additional stable I1 functions of thesens ors a re given in table V. These attitude se nso rs could be tested in the spa cecra ftby loosening a hexagonal nut (fig. 3) and pivoting the sen so rs to e ffec t a change in atti-tude signal.

    TEST E QU IP ME N T D E S IGNTo meet the crit ica l schedule require ments fo r the MCP qualification prog ramand delivery, the con tractor built three types of t es t equipment.1. Manufacturing test sets2. Manual test equipment (MTE)3 . Factory te st equipment (FTE)

    The manufacturing tes t sets were essential in the test activity associ ated with the pro-duction and ass embly of the MCP control as semblie s and prin ted wire networks. Thesetest sets performed satisfactorily and supported the program in a timely manner.The manual bench test console verified the operational s tat us of each redundantThis benchomponent in the M C P during acceptance tests and othe r box-level tests.console required that the input signal be switched manually at the time s req uired by thetes t specifications. Groups of t es t points (e. g., 60 test points) were collectively mon-itored and, if no anomaly occurred, that test zone of redundant elements within the M C Pwas considered satisfactory. This MTE was simil a r to the equipment developed fo r thecontrol pro gra me r in spa cec raf t 009 (AS- 201 mission) and was completed on Novem-ber 14, 1965, in time to support the initi al breadboard and ea rl y MCP prototypedeliveries.The MTE requir ed approximately thre e ti me s as long to complete a test run asdid the automatic FTE; however, the MTE was sufficiently simple that the equipmentcould be certi fied and debugged in a timely manner and could be reconfigured for com-patibility with changes in the flight hardware .The automati c FTE w a s used for the postenvironmental functional tes ts and the

    MCP syst ems tes ts . This te st equipment included the following.1. A punched-tape rea der to provide t he input st imuli with the a ssoc iate d powersupplies and signal- conditioning equipment2. A se ri es of interna l logic cir cuit s to control the switching and route the sig-nals to the correct M C P area


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    3. A ma st er clock to cont rol the timing of the input signals and to provide a t imecompari son of the MCP response4. An output load simulator to simulat e interfacing sys tem s loads5. A print er to provide a tape record of the test eventsThe automat ic FTE was pri maril y used to support qualification testing. A seriesof test tapes was pre par ed to support the environmental and postenvironmental func-tional tests as follows.1. Environmental functional te st s

    a. Abbreviated- time simula ted missionb. Real- time simulated missionc. Simulated abort and entry test

    2. Postenvironmental functional te st sa. SCC functional te stb. GCC functional te st

    The re quirement exis ted to automate the test and sequencing of the MCP while itwas operating in the qualification- te st environment. Fo r example, during the vacuumtest, an abbreviated-time simulated mission was performe d while the MCP was in the4-hour soak perio d of the vacuum environment. This requirement would have beenimpossible to achieve with the MTE because 48 hours would be required to sequence theMCP through all the pro gra me r functions manually. As a result, the requirement for a4-hour vacuum soak would have to be exceeded. However, if the 4-hour vacuum soakwere re tained, the number of functions that could be manually sequenced would be solimited that only a sm al l pa rt of t he MCP internal logic circu its could be tested.

    The development and certif ica tion of the elaborate and complex FTE within theallotted schedule period caused considerable difficulty. A 4-month delay in the startof qualification of the GCC and SCC was generally at tributed to proble ms in cert ifyingthe test equipment, the te st specifications, and the tes t tapes. The schedule problemconcerning certif ica tion of the FTE was rela ted to the original design concept and theplanned method of test. Considering the crit ica l development schedules and the sma llnumber of unmanned sy st em s to be delivered, the te st equipment concept was much toocomplex and automation was overly emphasized.

    Some speci fic problem areas in certifying the FT E included the following.1. The test equipment did not ve rify functional paths within the MCP but checkedout zones or groups of components; the refor e, when a hardware change was incorporatedin the MCP, a compatible change was difficult to incorporat e in the rela ted componentgroup of the test equipment. Hardware changes al so caused difficulty in updating thetest specifications.


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    2. The tape r ea de r had no reliable method of per forming an internal verificat ionor self-check. If a pa rt of the tape messa ge w a s missed, it was difficult to determ inewhether the problem was in the reade r, the test equipment, the M C P , or some othersystem or component.3. A reliable method was not developed to revise only specific sections of thetest tapes to reflect hardware modifications. A reprogram ing effort involving the enti retes t sequence seem ed to be required. The te st tape could not be cut and spliced ; ther e-fore, a new tape had to be generated to include the updated test section. A s a result,the manpower requir ements f or test equipment program ing we re increase d wheneverthe flight hardware was changed.Because of these tes t equipment repr ogra ming delays, the FTE was not used toany grea t extent in supporting the M C P reacceptance tests following design modifica-tions. The MCP redundancy te st perf ormed in the spacecra ft provided a sufficientconfidence level, and a systems -level functional acceptance te st at the vendor was notrequired. The FTE was not re program ed and reconfigured to reflect the numerousMCP hardware changes. The engineering time was mor e efficiently used in actuallyperforming the vendor box-level acceptance tes ts on the slower MTE than in pre pa rin gthe automatic FTE to perform the MCP system-level t est.This experience could well be applied to the development of t es t equipment forfuture programs that have smal l quantities of deliverab le end-item s.program, it seems preferable to expend the necessary engineering manpower in devel-oping simple, flexible, manual, general-purpose test equipment and then to make thenecessary allowances in delivery schedules. This approach appe ars pref erab le toexpending the manpower in developing automated, complex, inflexible te st equipmentthat would perform the test faster.

    Fo r this type

    DEVELOPMENT SCHEDULES AND TEST P R O G R A MOn June 25, 1964, the Apollo prime contractor w a s notified to develop a programerwith the capability to conduct the unmanned missions AS- 201, AS- 02, AS- 501, andAS-502. The original schedule for the MCP installation into spacecraft 011 at the con-tractor's facility w a s January 13, 1966. The MCP development team had 19 months todesign, build, test, and deliver the first flight syst em. The following paragraphsdescribe the most significant milestones concerning this development.


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    B r e a d b o a rd a n d P r o t o ty p e D e v e l o p m e n tThe following schedule w as achieved and indicates the co mpre ssed and cri tica lnatu re of delivery milestones fo r the MCP.

    Delivery milestone DateDesign configuration fre eze October 28, 1965Breadboard system deliveryFirst prototype unit deliverySecond prototype unit delivery

    November 1965December 3, 1965December 17, 1965

    First production unit delivery Janua ry 14, 1966Although the design configuration freeze was dated October 28, the following sig-nificant changes t o the MCP design were approved on November 8; therefore, theconfiguration was not really frozen.1. The on and off t im es of the flight-qualification tape re co rd er s we re changedand requ ired wiring changes in the MCP.2. The very-high-frequency antenna was switched differently for spacecraf t 017and 020, and additional wiring changes were required.A maximum of 1 month was scheduled between delivery of t he breadboard and thefirst prototype. The te rm "breadboard" cannot be used in the sen se that the breadboardwas a device to be tes ted and evaluated, with the re su lt s of the evaluations being fedback as design improvements. The rigo rous acceptance te st s and inspection-approval

    criteria that normally constra in development did not apply to thi s breadboard unit;ther efor e, the manufacture r could produce the unit as a working device to help in thete st equipment development and certificat ion. The prototype unit used the s am e produc-tion manufacturing and assembly techniques as the flight units. The f ir s t prototype wasdelivered to the spa cecraf t contra ctor for simulation testing and interfa ce verificationtesting. The combined sys tems te st s and simulations, using the first prototype unit,uncovered the problem of the MCP rel ay drive rs triggering on noise. These evaluationte st s were a lso valuable in establishing a redundancy checkout schem e fo r the MCPwhile it was installed in the spacecraft.The second prototype was used as a prequalification te st a rti cle fo r certifyingboth the MTE and the FTE before the official st ar t of the qualification progra m. Suffi-

    cient time was not available fo r the breadboard- and prototype-development pro gra msto provide useful information fo r the flight system design without a significant cost andschedule impact. Ideally, 6 months should be scheduled between the breadboard andfirst production item delivery dates f o r hardware as complex as the MCP.E l e c t r o m a g n e t i c - In t e r f e r e n c e C o n s i d e r a ti o n s

    While evaluating the first MCP prototype in the communications laboratory , the- a errnft- - - -- - rnntmrtnr- --.. .. discovered that the GCC relay dr iv er s were triggering on noise33

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    voltages. At the beginning of the Apollo Pr og ra m, one MCP unit was scheduled fo renvironmental qualification tests and another unit for em i qualification tests.test and success values were difficult to es tabl ish on a black-box level because theinterference is an interrelated- syst ems problem.qualification tests on the black-box level was eliminated before the MCP qualificationtests were scheduled. An overall em i test scheme was to be establishe d on thespacecraft- test level.

    The emiTherefore, the requirement fo r em i

    Te s t Equ i me n CertificationThe following schedule was achieved concerning certification of the FTE.

    Item DateFTE test tape development start October 28, 1965MTE completion November 14, 1965FTE test tape completion March 7, 1966MTE recalibration March 11, 1966FTE test tape certification Apr il 6, 1966FTE certification April 8, 1966Certification of the FTE was important in that th is certification was a constraintto the st ar t of the sys te ms tes t portion of the MCP qualification pro gra m. The FT Ecertification, o r development test ing, could not begin without an MCP te st article toproce ss the response s to the test input signals.in November 1965, was used in th is development.6 months for development and certifica tion testi ng of the FTE.

    was extremely shor t fo r testing, debugging, and certifying a test equipment system ofthi s complexity. However, the orig inal schedules allowed only a 2-month period frombreadboard del ivery to cert ification completion and qualification te st start.period was not sufficient to achieve the test equipment certificat ion; the refo re, the quali-fication st ar t date was extended by 4 months.

    The MCP breadboard syst em, deliveredThis late delivery allowed onlyThis length of t im e


    The cer tificat ion of the test equipment was achieved by using a production proto-type MCP unit that was essenti ally identica l to the qualification unit t o be tested later.First, an acceptance test using manual test methods was performe d on the prototype,and each redundant function was verified to be operating. This unit was then used as atest equipment certification unit, and the s am e test specifications wer e used. If everytest function was proc ess ed through the ce rtification unit and was recorded by the F TEwith no anomalies, the test function was certified. If an anomaly occurred, then ananalysis had to be performed t o determine whether the test equipment or the certifica-tion unit had malfunctioned. Thi s step-by- st ep method wa s demanding and ti me con-suming, but the FTE was finally certified.


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    P r o d u c t i o n D e li ve r yThe following schedule w a s achieved f or the M C P production unit de live ries andmodifications.

    Production delive ry DateUnit 1Unit 2Unit 3Completion of design modifications to unit 1 afterUnit 4Modification of unit 3 to spacecraft 017 configurationUnit 5Unit 6Modification of uni t 4 to spacec ra ft 01 7 configuration

    completion of contract or te st s

    Janua ry 14, 1966March 30, 1966April 5, 1966May 1966June 11, 1966July 14, 1966August 19, 1966September 16, 1966October 3, 1966

    Of the six production units delivered, units 1, 3, and 4 required sev era l designmodifications to make them compatible with the M C P design configuration for space-craft 017 and 020. The design changes were incorporated in production units 5 and 6before delivery. The qualification unit 2 did not require modification because the designmodifications did not requi re requalification testing. The MCP supported the spacecraftdelivery and test schedule dates; however, some of the des ign changes and rework hadto be accomplished during the idle vehicle test periods. For example, the rework wassta rted aft er the MCP finished supporting t h e integrated sy ste ms tes t at the spacecraftcontractor's facility and w as completed before the next requirement to support tests inthe vehicle at the NASA John F. Kennedy Space Center (KSC). Ideally, the spacecraf tconnectors would not have been disturbed, and the MCP would have been delive red tothe KSC while installed in the spacecraft.

    Q u a l i f i c a t i o n T e s t sThe following key schedule dates describe the qualification te st progra m.

    Item DateOriginal qualification test startActual qualification test startMTE certificationQualification production unit 2 deliveryQualification production unit 3 deliveryFTE certif its. inn

    December 17, 1965Febru ary 17, 1966March 11, 1966March 30, 1966April 5, 1966April 8, 1966


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    Item DateQualification test completionQualification report release

    April 23, 1966May 31, 1966

    The orig inal qualification te st start date is listed to emphasize the importance ofallowing adequate time to certif y and evaluate te st equipment. Thi s orig ina l start dateof December 17, 1965, w as postponed 4 months for the GCC and SCC units because ofpreviously mentioned problem s with te st equipment certification. The actu al qualifica-tion program was able t o be begun as ear ly a's February only because the qualificationtesting of the ADS unit w a s started before the GCC and SCC unit's. The FTE w a s notrequired f or the postenvironmental tests of the ADS unit. The FT E was finally certifiedApril 8, 1966, and was available for u se during the MCP system-level functional te st sfor postenvironmental evaluations. These system t es ts of th e MCP were program ed onpunched tape, and the test equipment automatically generated, switched, and routed thestimulus and response signals; mea sured the ti me of response; and evaluated the logicst at e of the cir cu it ry being tested.

    A detailed schedule of the qualification tes tin g sequence is shown in figu re 13.Items 1 o 15 n figure 13 repr esen t keywords f or coding the test activity during anyspecific tes t period. Fo r example, f rom March 10 o March 15, during the qualificationte st of sys tem 2, the activity w as 15 (MTE functional test s). As the tes t resu lts arediscussed in the following para grap hs, the test sequence can be established by referringto figure 13.

    Qualification testing,system 1

    Ground command controllerand spacecraft commandcontrollerQualification testing,system 2Spacecrafl commandcon ro I erGround commandcontrol ler

    1. Vacuum 8. Acceleration2. Oxidation 9. Factory test equipm ent test3. Humidi ty 10. Manufac tur ing rework4. Resistance 11. Manua l test equipment test and specification certification5. Random vibration 12. Factory test equipme nt test and specification certif icatio n6. High temperature 13. Life test vibratio n7 . Shock 14. Real-time mission simulated mission

    15. Manual test equipment tests functional

    m4111213[ 11 I 15 I 12 IlOld 12 I 13 I 12 ] mttitude and L 1 1deceleration sensorI 1 I I I I 1 1 I I I I I 1

    February March Apr i17 20 25 ' 1 5 10 1 5 20 25 I 1 5 10 15 20 25

    Figure 13.- Qualification testin g sequence.The ADS package of MCP production unit 2 was subjected to the qualification tes t

    Because this package als o hadenvironments.mental tests, except shock, were completed in 1week.Because of the simple design of th is sen sor package, all the environ-


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    to support the life test , the shock environmental test was postponed until the end ofqualification testing. The sensor package successfully completed the qualification testswith no anomalies, no visible physical damage, and no operational degradation,The MCP production unit 2 (qualification test unit 2) w a s initially ready to beginqualification tes ting on Fe brua ry 26, 1966. However, the FTE ei ther had not been

    completed or was not certifi ed. Because of the crit ical schedule requiremen ts, thequalification tests were star ted, using only a few manually initiated commands fo r eachpostenvironmental ver ification of the MCP. No failur es were detected during the initialvacuum, oxidation, o r humidity testing .After the humidity test, the qualification test packages were to be given a completepackage functional test , using the MTE.plete the te stin g of qualification unit 2, but the new tes t spec ifications re qui red box-leveltest ing of redundant circuitry. These specifications had not been checked out againstany package o r with the MTE. When the postenvironmental (vacuum, oxidation, andhumidity) te stin g was attempted, numerous problems were encountered and too many

    unknowns (such as MTE, te st specif ications, and MCP) were involved. As a result, theperiod between March 2 and April 2, 1966, was used to debug and certi fy the specifica-tions and the MTE, to re te st the MCP, and to check out the functional test tapes for theFTE. During thi s period, fai lure s were detected in the packages; s ome failur es weredue to manufacturing er r o r s not previously tested in the redundant c irc uits during sell-off; others were induced by the MTE. These failures reemphasized the critical require-ment of enter ing a qualification test program with a good baseline; that is, with certifie dte st equipment, verifi ed proc edures, and adequate specifications.

    This equipment had been used earlier to com-

    After qualification unit 2 finished servi ng as a test arti cle fo r the certification ofthe automatic FTE, the tes t equipment was successfully used to complete a functionaltest on qualification unit 2. The unit then entered the life test sequence on April 2, 1966.The purpose of th is test was to verify that the M C P could perform normal missionfunctions aft er accumulating mo re than 500 hours of operating time. A f t e r the requirednumber of operating hours was accrued , the M C P entered the real-t ime simulatedmission run on April 15, 1966. The MCP proved to be capable of per forming the func-tion of a rea l-ti me mission af ter being subjected to random vibration levels and accruingmore than 500 hou rs of operating time.

    The MCP production unit 3 (qualification test unit 1) entered the vibration environ-ment portion of the qualification te st on April 2, 1966.investigated wer e the resonance s of each package (resonance se arc h) and the suscept i-bility of the MCP to random vibrat ion. The postvibration phys ical inspection of the MCPindicated 35 inst ance s of fra ct ured sol der joints on the pins of the SCC contro l assembl yconnector boar ds (fig. 14). The GCC had 10 loose or broken sold er joints aroundsi mi la r pins. However, no functional tes t failur es were attributed to the solder frac-tu res around the pins.included solder ing the termin als on both sid es of the ci rcui t st ri p (fig. 15) and addinga bracket to improve the wire- bundle routing. The corre ctiv e action was successful,and the problem did not re cu r during future vibration tests.

    The vibration effects to be

    The corrective action f o r the cracked- solder- joint problem


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    (a) Example 1.

    (b) Example 2.


    (c) Example 3 .Figure 14. Fractured solder terminals.

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    (a) Orientation of rework area. (b) Closeup of rework area.Figure 15. - Repair method, solder ing backside of terminal str ip .

    The second significant problem was detected during the humidity and posthumidityfunctional test portions of the qualification program. The re ve rs e impedance of theMCP diode quads was below the specification limits after the 16-hour humidity test.These impedances were within the specified value (grea te r than 700 kilohms) after about1 hour of drying. The diodes in the SCC were affected after being exposed to95 + 5 percent rel ative humidity; but, for normal unmanned flight, humidity was notexpected to be a problem. The cor rec tiv e action was to provide added protection byapplying polyurethane (polycoat) to the control assemblies containing the diode quads(fig. 16). On April 22, 1966, af te r the polyurethane was applied to the control assem-blies in qualification unit 2, the unit was retested in the humidity environment andsatisfactor ily met the specifications. The qualification tests were completed onApril 23, 1966, and the test report was released on May 31, 1966.

    I n t e r f a c e V e r i f i c a t i o n TestsThe inter face verification tests performed in the various engineering laboratori esat the spacecraft contractor 's facil ity provided much useful data. Some of the mostsignificant re sul ts were as follows.1. Identification of the relay dr iv er emi problem in the GCC2. Establishment of the concept of onboard redundancy tests for the MCP3. Verification of the allowable SPS gimbal position mistr im pa rame te rs4. Verification of the new design modifications before actual installation


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    5. Identification of the system int erface incompatibil ities6. Provision of useful information for resolving spacecraft test anomalies

    Figure 16. - Typical diode mounting bracket.The MCP prototype unit 1was delivered to the spacecraft contractor on Decem-ber 3, 1965. This unit was first checked to verify electrical interface and compatibilitywith the spacecraft electri cal power system. Unit 1 w a s then subjected to s eve raldifferent interface tests with individual spacecraft systems, such as the communications

    sys tem and the MESC. Finally, the M C P prototype was tested, along with severalother systems, in the guidance and control labora tory during the combined systemsdynamic verification test s. The combined syste ms-te st setup and the use of toggleswitches to switch the internal grounds (G1, G2, and G3) of the MCP suggested a methodfo r testing redundancy of the M C P while in the spacecraft.


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    Spacecraft TestsThe M C P w a s installed in the spacecraft before the combined and integrated sys-tems tests were star ted at the spacecraft contractor's facility. The MCP served as theinterfacing unit between the G&N system and other input stimulus s our ce s and the space-cra ft flight sy ste ms that actually activated and performed the d esir ed s pacec raft output

    functions during the ground tes ts and also during flight. The MCP had no flight orspac ecraft tes t measur eme nt allocation, even though it had 10 connecto rs with over500 measure ment pins readily available for bench te st s at the factory.for having no MCP flight meas urem ents nor ground test mea sure men ts was that the un-manned vehicles would be instrumented and tested the same as the manned vehicles.Therefore, the functional operation of the M C P was de termine d by observing the func-tional operation of the re late d output sys tems that were instrumented. This rationalewould have been adequate if the MCP process ed progra mer signals through single func-tional paths. However, the MCP contained numerous ser ie s- redundant and para lle l-redundant path s (as previously described).

    The rationale

    The Apollo Program had a requirement that each redundant path be verified asfunctioning properly j ust before launch. This requirement was interpreted to mean thatthe MCP, even though it was for unmanned flights, had to have its redundant paths veri-fied in the s pacecr aft ju st before launch. The following schedule indicates the timerequired to implement the MCP redundancy tests.Event

    NASA direct ed the cont rac tor to accomplish space-NASA management met with the contractor to

    craft redundancy tests.

    resolve details concerning spacec raft redundancycheckout requirem ents.ment me asur eme nts fo r fault isolation of the M C P .Contractor requested 78 automat ic checkout equip-

    Measurement requirement request w a s denied.NASA review dete rmine d that no plan w a s availableMCP spacec raft redundancy tes t plan was initiated.Decision was made not to verify the MCP redundancy

    with acceptance checkout equipment.The contractor proces sed an internal procedure tover ify MCP redundancy.An NASA management official directive emphasizedrequirem ent to perf orm MCP redundancy at theKSC for spacecraft 011, 017, and 020, and at thecon tra cto r's facilit y fo r spacec raft 017, and 020.

    f o r installed MCP redundancy test.

    The first MCP redundancy test was performed atths F3C 02 EpaCecr2ft 011.

    DateMarch 25, 1965April 1965

    Apri l 20, 1965October 1, 1965November 1965December 1965February 3, 1966April 21, 1966June 9, 1966

    July 14, 1966


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    During the l- ye ar period between the original directive and the final processing ofprocedures for performing this test, the co ntra ctor maintained that the MCP redundancytests would not be advantageous fo r the following r ea so ns .1. A 50-man-month effort in prog raming co st fo r acceptance checkout equipmentcould be saved.2. A saving of 120 hours of spacecraft tes t time would resul t, compared with the12 hours required f or bench te st equipment.3 . Inte rface equipment fo r the acceptance checkout equipment would have to bedesigned, fabricated, and certified.4. The acceptance- checkout- equi