Gemini Spacecraft Electrical System Specification

8/7/2019 Gemini Spacecraft Electrical System Specification

1/68

& DATE 20 F E B R U A R Y 1962REVISED 15 NOVEMBER 1962

G E M IN I S P A C E C R A F TE L E C T R I C A L S Y S T E M S P E C I F I C A T I O N [ U )

REPORT 8638 SERIAL NO.

CONTROL NO. c~67466

MCDCHVIMELLARCRAFT CO*OFATO

N T C O N T A I N S I N F O R M A T I O N A F F E C T I N G T H E N A T I O N A L D E F E N S EL A W S . TITLEi 794, TH E TRANSMISSION ORNNER TO AN U N A U T H O R I Z E D P E R S O N i s PROHIB ITED G Y

SUBMITTED UNDER N A S A C O N T R A C T NAS 9-170

GEHUI- - - .' ' -- ' >Unclas


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D A T E 2 0 February 1962R E V I S E D 6 June 1962RE V I S E D ,

MOOOA/A/HST. LOUIS, MISSOURI P A G E

R E P O R T .M O DE L .

8638Gemini

Prepared by:

Approved by:

APPROVALS

F b. Jcmes - Specifications Group

L. M. Parker - Sr. Specifications Engineer

G. J< keber - Assistant Project, Engineer hlectrical

W. J. Bmtz - ProjectTEngineer


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D A T E 20 February 1962REV^SEi 11 JUIV-IffST'; '"""R E V I S E D 15 November 1962

MCDONNELLST. LOUIS, MISSOURI PAGE

MODEL Gemini'

PARAGRAPH

1.1.12.2.12.22.32.3.12.3.22.3.32.U3.3.13.23.33.3.13.3.1.13.3.1.23.3.1.33.3.1.U3.3.1.53.3.1.63.3.1.73.3.1.83.3.1.93.3.23.3.2.13.3.2.23.3.2.33.3.2.U3.3.33.3.3.13.3.3.23.3.3.33.3.3.U3.3.3.5

TABLE OF CONTENTSTITLE

TITLE PACEAPPROVALSLIST OF TABLES AND FIGURESLIST OF EFFECTIVE PAGESINDEX OF REVISIONSGENERALScopeAPPLICABLE DOCUMENTSGeneralGovernment DocumentsMcDonnell DocumentsReportsDrawingsProcess SpecificationsSupersedenceDESCRIPTIONGeneralPower DistributionSystem ManagementVisual IndicationsPower Transfer Warning LightBattery Condition IndicatorAmmeterVoltmeter2 Quantity and PressureH2 Quantity and PressureStack Malfunction IndicationCoolant Pump Warning LightsFuel Cell AmmeterSwitchesSwitches - Fuel-Cell Battery SubsystemSwitches - Silver- Zinc Battery SubsystemBus Tie SwitchVoltmeter Selector SwitchMission Electrical ProceduresPre launchLaunch Through InsertionCatch-Up Through OrbitRetrograde and Re-entryLanding and Post-Landing

PAGE NO.

iivVvi11

1 111112U66677777101010101010

1010121212121212131313


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DAT E 20 February 1962 MCDONNELLST. LOUIS. MISSOURIR E V I S E D 6 June 1962REVISED November 1962

P A G EREPORT.MODEL .

iii8638

Gemini

PARAGRAPH3.U3.53.5.13.5.23.5.2.13.5.2.23.5.33.5.1*3. 5. U.l3.5.U.23.5.U.33.5.U.U3.5.53.5.63.5.73.63.7U.U.I5.5.16.6.17.7.18.8.18.1.18.1.38.1.U9.9.1

Appendix IAppendix IIAddendum A

TABLE OF CONTENTS (Continued)TITLE

Circuit ProtectionComponentsRequirementsBatteriesFuel-Cell Battery SubsystemSilver-Zinc BatteriesRelaysSwitchesToggle SwitchesRotary SwitchesLimit SwitchesTelelights and Telelight/SwitchesExternal Power ConnectionsConnectorsCircuit BreakersFabricationInstallationENVIRONMENTAL CONDITIONSGeneralQUALITY ASSURANCEGeneralRELIABILITYGeneralDATA REQUIREMENTSGeneralTESTINGGeneralBattery TestsSpacecraft Systems Test (SST)Simulated Orbital MissionDEFINITIONSGeneralAPPENDICESEQUIPMENT LISTCHANGE LISTELECTRICAL SYSTEM - SPACECRAFT NO. 1

PAGE NO.Ik1515151518181919191919192020202122222222222222222323232k2k2525

3739Al


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DATE 20 February 1962 MCDOhlhlEST. LOUIS, MISSOURI PAGE ivBcrx/icrn

TABLETable I

Table II

Table III

FIGUREFigure 1Figure 2Figure 3Figure hFigure 5

Figure 6

Figure 7

Figure 8

Figure 9

Figure 10

Figure 11

Figure 12

Figure 13

LIST OF TABLESAND FIGURES

TITLEEnvironmental Conditions for Electrical EquipmentLocated in the Pressurized CabinEnvironmental Conditions for Equipment Located inthe Re -Entry Module External to the PressurizedCabinEnvironmental Conditions for Electrical EquipmentLocated in the Adapter

Installation - Electrical Power System (illustration)Electrical Power System (Schematic Diagram)Power Management Controls and Displays (illustration)Fuel-Cell Battery Subsystem (Schematic Diagram)Stack Assembly - Fuel-Cell Battery Subsystem(illustration)Gemini Spacecraft Vibration Spectra EquipmentCategory AGemini Spacecraft Vibration Spectra EquipmentCategory BGemini Spacecraft Random Vibration PowerSpectral DensityGemini Spacecraft Acoustic Environment Spectrum(Re -Entry Module - Inside Pressurized Cabin)Gemini Spacecraft Acoustic Environment Spectrum(Re-Entry Module - External to Pressurized Cabin)Gemini Spacecraft Acoustic Environment Spectrum(Adapter)Gemini Spacecraft Acoustic Environment Spectrum(External or Partially External to the Spacecraftprimary structure)Gemin , Spacecraft Design Load Factors

Gemini

PAGE

26

27

28

58916

1729303132333 * "

3536


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DATE 6 June 1962 MCDONNELLST. LOUIS, MISSOURI PAGEREVISEDREVISED

November 1962 REPORT.MODEL .

8638Gemini

LIST OF EFFECTIVE PAGESThe pages of this report currently in effect are listed below in numerical order.TITLE PAGEi through vi1 through U3Al through AID

REV


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DATE 6 June 1962 MOOOJVJVEST. LOUIS. MISSOURIREVISEDREVISED

July 1962November 1962

PAGE .REPORT 868MODEL Gemini.

INDEX OF REVISIONS

DATE

6/6/62

6/6/62

7/11/62

11/15/62

PAGES AFFECTEDREVISEDTitlePageiii - iv1 - 35 - U*ii and

vi6, 10,11, 18,37, 39and UO1 - 1 21U, 15,16, 18,19, 20Title Pg,vi

ADDEDv & vi1 5 - l a

1*2, 1* 3Al - A10

REMOVED REMARKSSee Appendix II

See Appendix II

See Appendix II

REVISED BY

F. E. Jones

F. E. Jones

R.D. Korando

APPROVED

L.M. Parker

L.M. Parker -

L.M. Parker


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D A T E 20 February 1962R E V I S E D 6 June 1962R E V I S E D I1? November 1962

MCDONNELLST. LOUIS, MISSOURI P AG E

REPORT.MODEL .

8638Gemini

1.1.1

GENERALSCOPE.- The purpose of this specification is to delineate theapplicable documents, description, environmental conditions,reliability and quality assurance provisions, data requirements, testing, anddefinitions pertinent to the Gemini Model 133? spacecraft electrical system.This document is not to be construed as, or used for, an inspection guide.

2.2.1

APPLICABLE DOCUMENTSGENERAL.- It is the contractor's intent, relevant to the use ofGovernment specifications in the design, fabrication and instal-lation of the electrical system, to utilize existing specifications wherepracticable. In cases where the subject matter is applicable but the specificrequirements are not compatible due to the advanced design of this system, thespecifications and documents referenced are followed only to the extent thatthe intent of such requirements aremet.

2.2 GOVERNMENT DOCUMENTS.- Government specifications, standards, andpublications listed in McDonnell Report 8357 form a part of thissystem specification to the extent specified in Paragraph 2.1.2.3 MCDONNELL DOCUMENTS.- The following McDonnell documents areapplicable to this system specification insofar as their contentconcerns the system described herein.2.3.1

2.3.2

REPORTS83578518

8580-38580-58580-7

8580-88612/MB-10U8DRAWINGS52-7970052-79702

"Gemini Spacecraft ApplicableDocuments""Electrical and Electronic Equip-ment Design Requirements forModel 133?""Project Gemini Reliability Plan""Project Gemini Test Plan""Quality Assurance Provisions(Plan) for Project Gemini SpaceSystem""Model 133P Gemini ProgramDocumentation Plan""Gemini Spacecraft/Launch VehicleInterface Specification"

Battery, Fuel CellBattery, Storage


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MCDONNELLR E V I S E DREVISED

2.3

2.3

20 February 19626 June 1962

15 November 1962

.2 DRAWINGS52-7970352-7970552-7970652-7970852-79710

52-7971152-7971352-7971552-7971752-7972152-7972252-797252-8170752-8370152-90000

.3 PROCESS1333^

16001.51700

171*1017^10.1

ST.LOUIS, MISSOURI PAGE 2REPORT 8638

- (Continued)Relays, Control and PowerSwitches, ControlRotary SwitchGround UnbilicalTelelights and Telelight SwitchesFloodlightPower RelayRelay, SpecialShunt Resistor AssemblyCircuit BreakerDiodeLight Switch AssemblyVoltmeterReactants Supply SystemFinish Specification

SPECIFICATIONS. -Preparation and Application of Coating toInterior Surfaces of Sealed Cabin Area ofModel 133Marking of Model 133 Parts and AssembliesInstallation of Electrical Wiring in Model133Fabrication of Electrical Wire Assembliesfor Model 133Assembly of Electrical Cable Terminals andSplices for Model 133

REV


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DATE 20 February 10,626 June 1962

R E V I S E D 15 November 1962

ST. LOUIS, MISSOURI P A G ER EV I S ED REPORT.

MODEL .868

2.3.3 PROCESS SPECIFICATIONS - (Continued)17U10.2

20106

20500

Assembly of Electrical Connectors forModel 133Storage and Handling of Silver-ZincBatteries for Model 133Fabrication and Housekeeping PoliciesApplicable to Model 133

RV


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MCDONNELLST. LOUIS, MISSOURI P A G E

REVISEDREPORT.MODEL .

8638Gemini

2.U SUPERSEDENCE.- If any of the specifications, standards, drawingsand publications which form a part of this system specificationare superseded during the life of the contract of which this system specifica-tion is a part, the later issue may be used.


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D A T E 20 February 1962' July 1962

MODOJVJVST. LOUIS, MISSOURI PAGE

R E V . S E DREVISED 15 November 1962

REPORT.MODEL .

8638Gemini

3.3.1

DESCRIPTIONGENERAL.- The function of the Gemini spacecraft electrical systemis to supply and distribute electrical power of the proper voltageto all spacecraft devices requiring electrical power for operation, with aminimum of interference between the various subsystems and devices. Powersources are capable of supplying sufficient power to the electrically-operatedequipment for completion of a lii-day orbital mission or a 2-day mission inclu-ding rendezvous and docking, plus a post-landing period of 12 hours and a pre-launch period of two hours. Power consumption is minimized to the greatestpracticable extent compatible with system requirements.

3.1.1 P010ER SOURCE.- The electrical system is a 2-wire, grounded systemutilizing fuel-cell and silver-zinc batteries as described in Para-graph 3-5.2 and its subparagraphs as sources of DC power. The fuel-cell batterysubsystem is located in the equipment section of the adapter. The silver-zincbatteries are located in the re-entry module of the spacecraft, outside thepressurized area.3.1.2 AC POWER.- Devices utilizing AC power are supplied by self-con-tained inverters within the individual systems. These invertersare not considered a part of the electrical system, but are treated as inherentelements of the system which they serve.3.1.3 WIRING AND CONNECTORS.- For purposes of this specification, the

wiring and connectors which interconnect components of systemsother than the electrical system are considered a part of the electrical systemand meet all applicable requirements specified herein.3.2 POfeER DISTRIBUTION.- DC power distribution to the utilizing systemsis accomplished through a main bus subsystem and an isolated bussubsystem each of which is independent of the other. (See Figure 2 for powerdistribution schematic).3.2.1 MAIN BUS SUBSYSTEM.- In a normal mission, the main bus is energizedby the fuel-cell batteries from prelaunch until approximatelyTr-AT minutes, at which time the main silver-zinc batteries are manuallyswitched to the main bus circuit. The silver-zinc batteries will be manuallyswitched to the main bus in parallel with the fuel-cell battery subsystem duringthe launch phase of the mission to insure continuity of electrical power incase of an abort.

REV


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MGDOJVJVfiST. LOUIS, MISSOURI PAGE

REPORT.MODEL .

86^8Gemini

3.2.2 ISOLATED BUS SUBSYSTEM.- The isolated bus subsystem consists oftwo squib buses and a control bus energized by three isolatedsilver-zinc batteries throughout the mission. The functions of this subsystemare to supply power to the pyrotechnic ignition devices through redundant,isolated, parallel circuits from power source to igniter and to supply power tosolenoids and control relays. The control bus is energized by both squib busesand a separate battery, but isolated from each by diodes. In the interest oflaunch site and in-flight safety, provisions are incorporated for disarming thepyrotechnic circuits. A master arm switch is provided in the blockhouse.3.3 SYSTEM MANAGEMENT.- Provisions are incorporated on the instrumentpanel for monitoring and managing the power source. Visual in-dications and control switches are depicted in Figure 3 and described insucceeding paragraphs. Monitoring and switching procedures will be followed,consistent with readings obtained throughout the mission. A tentative programfor these operations is outlined in Paragraph 3-3-3.3.3.1 VISUAL INDICATIONS.- Visual indications of electrical system events,conditions, outputs, and malfunctions are provided, consisting ofwarning and indicator lights (see Paragraph 35-^*0 and single and multiple-reading meters.3.3.1.1 BATTERY POWER WARNING LIGHT.- At Tr-5minutes, an amber warninglight located in the sequence-light cluster along the left-handside of the center section of the instrument panel apprises the astronauts thatpower usage should be transferred from the fuel-cell source to the silver-zincsource due to impending separation of the adapter equipment section. When allmain battery switches (see Paragraph 332.2.l) are in the ON position, thelight is illuminated green.3-3.1.23-3.1.3

BATTERY CONDITION INDICATOR.- Deleted.SECTION AMMETER.- A dual-reading ammeter is installed on the righthand section of the instrument panel. The vertical scales covera range of 0 to 50 amperes. During the launch through insertion phase of themission the ammeter will indicate the current flow of: (l) Stacks 1A through

IB and silver-zinc batteries Ml and M2; (2) Stacks 2A through 2C and silver-zinc batteries M3 through M5. During the orbital phase fuel-cell sectioncurrents are indicated. During re-entry, landing and post landing the ammeterindicates the silver-zinc battery currents. (See section 3-3.3.)3.3.1. VOLTMETER.- A voltmeter and associated selector switch is providedon the right-hand section of the instrument panel. The selectorswitch provides a means of monitoring fuel-cell stack voltages, common controlbus voltage, squib bus 1 and 2 voltages and main bus voltage.


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DATEREVISEDREVISED

20 February 19&26 June 1962

1$ licwMbar 1962

MCDONNST. LOUIS, MISSOURI P A G E

R E P O R T .MODEL .

ELECTRICAL POWER SYSTEM SCHEMATICLOADS LOADS

MAIN BUS

3AGE-MAIN '

BUS-1SW ITCH GND. TESTP O W E R

JSQUIB BATT ON

cn

AG-ZNS-l

SW .

rr FFOHSQUIB BATT.S- l PWR SW.

I1I"

*io-M__^^ J ^~ ]

>c^x' 2o_


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.S T A C K. 2-B flS T A C K2- C

P U R G E _ S U P P L Y _ _ _)

I 1.)SECTION CONTROL SWITCH (ON)(A) OPENS H20 DRAIN(B) APPLIES H2 & Oj R E A C T A N T

2.) C O O L A N T INTERLOCK(A ) COOLANT PUMP MUST

BE O P E R A T I N G3.) SECT IO N POWER SWITCH (ON)

(A) CONNECTS SECTION TO BUSIF ITEM 1 AND 2 ARE SATISFIED!4.) S T A C K R E M O V A L SWITCH (OFF)

(A ) ELECTRICALLY DISCO NNECTSSTACK FROM BUS(B) SHUTS OFF H, TO(A ) EL"(B) SHIINIINDIVIDUAL STACK

I.) SILVER-ZINC BATT. CONTROL SW5.' (A) CLOSED DURING LAUNCH

(B) OPEN DURING ORBIT(C) CL O SED AT TR - AT

S A F E


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"Page missing from available version"


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D A T E 20 February 1Q62R E V I S E D 11 July 1962R E V I S E D 13 November 1962

JlfOOOJVJVCST. LOUIS, MISSOURI P A G E 10

TV ! 17 T\ T71 TV T T* TATREPORT.MODEL .

8638Gemini

3.3.1.5 02 QUANTITY AND PRESSURE.- A dual-reading, vertical-scale instru-ment is provided on the center section of the instrument panelfor monitoring the fuel-cell oxygen supply.3-3.1.6 H2 QUANTITY AND PRESSURE.- A dual-reading, vertical-scale instru-ment is provided on the center section of the instrument panelfor monitoring the fuel-cell hydrogen supply.3.3.1.7 STACK MALFUNCTION INDICATION.- A horizontal row of six telelight/switch assemblies is provided on the right-hand instrument panelfor indication and override of unsatisfactory conditions in individual fuel-cellstacks. An amber telelight indicates low stack current, no voltage, or reversecurrent operation. The switch portion of the telelight assembly disconnectsthe affected stack from the bus. (See Paragraph 3.3.2.1.5.) Illumination ofred lights in three of the six telelight/switch assemblies indicates an out-of-tolerance pressure differential in the corresponding fuel-cell battery section.3-3.1.8 COOLANT PUMP WARNING LIGHTS.- One warning light is provided foreach coolant pump in the redundant cooling loops. (These loopsare common to the fuel-cell, cabin, and equipment cooling systems and are pro-perly part of the environmental control system. The warning lights are refer-enced here only because they are included in the fuel-cell monitoring procedure(see also Paragraph 3.3.2.1.6). These lights are illuminated when an energizedcoolant pump motor becomes inoperative.3.3.1.9 FUEL-CELL AMMETER.- A fuel-cell ammeter is provided for measuringthe current flow through any stack. Selection of a particularstack is accomplished by a fuel-cell ammeter switch located on the right-handsection of the instrument panel (see Paragraph 3 3 2.lA ). The ammeter rangeis 0-20 amperes.

REV

NEW

3 - 3 . 23-3.2.13-3.2.1.1

SWITCHESSWITCHES - FUEL-CELL BATTERY SUBSYSTEMCONTROL SWITCHES.- Individual control switches are provided foreach fuel-cell battery section. These 2-position (ON-OFF) switchesare utilized to control reactant flow to, and by-product water from, the fuel-cell battery subsystem. Selection of the ON position of each switch openssolenoid valves which permit reactant and water flow, activating the section.This switch also acts in conjunction with the section power switch (see Para-graph 3.3.2.1.2) as a protective series connection which prevents applicationof an inactive fuel-cell section to the main bus.

3-3.2.1.2 POWER SWITCHES.- Individual 2-position (ON-OFF) power switchesare provided for application of each section of the activatedfuel-cell battery subsystem to the main bus.


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D A T E 20 February 1962 ST> LOUIS> M1SSOURI PAGE 11.R E V I S E D 11 July 1962 REPORT 86 8R E V I S E D 15 November 1962 G&&&&&&tflt4lr M O D E L Gemini

3.3.2.1.3 PURGE SWITCHES.- Individual 3-position (H2-OFF-02) momentary con-tact switches are provided for purging each section of the fuel-cell battery subsystem. Selection of the H2 position of the switch openssolenoid valves in the hydrogen circuit downstream from the battery section,permitting expulsion of any impurities within the section. Selection of the62 position performs the same function for the oxygen circuit. Immediatelyupon release of the actuated toggle, the switch returns to the OFF position.3.3.2.1.14. BCI SELECTOR SWITCH.- Deleted.3.3.2.1.5 STACK DISCONNECT SWITCHES.- Individual switches are provided forwithdrawing any desired fuel-cell stack from the main bus circuitin case of stack malfunction or non-operation. These switches are integralparts of the telelight/switch assemblies used for stack malfunction indication(see Paragraph 3.3.1.7).3.3.2.1.6 COOLANT PUMP SELECTOR SWITCHES.- Individual 2-position (ON-OFF)selector switches are provided for selective operation of thecoolant pumps in both primary and secondary cooling loops. (These switcheslike the associated warning lights, paragraph 3>3l'8j are part of the environ-mental control system and are referenced here because of inclusion in the fuel-cell monitoring procedure.)3.3.2.1.7 HEATER SWITCHES.- Individual 3-position (AUTO-OFF-MAN) switches areprovided for the fuel-cell hydrogen and oxygen supplies. Theseswitches are mounted on the center section of the instrument panel immediatelybelow the fuel-cell reactant quantity and pressure indicators, and are employedto control heating elements mounted in the reactant storage tanks. In theAUTO position, power to the elements is controlled by a pressure switch whichcloses when pressure drops to a preset level. As the reactant temperature rises,pressure increases until the pressure switch reaches its cutoff point, at whichtime the heating elements are de-energized. In the MAN position the 09 and/orH2 reactant pressure gauges are monitored for an optimum pressure level set forthin M.A.C. SCD 52-83701. The MAN position is momentary, therefore upon release,the power is removed from the heating element(s).3.3.2.1.8 LOGIC OVERRIDE SWITCH.- Deleted.3.3.2.1.9 CROSS-FEED SWITCH.- A 2-position (X-FEED - NORMAL) switch is pro-vided for controlling two normally-closed solenoid valves locateddownstream from the pressure regulators in the fuel-cell reactant supply lines.Selecting X-FEED position energizes the solenoid valves, enabling reactantsfrom one pressure regulator to enter the supply lines normally served by theother regulator. This permits use of a fuel-cell section which has been de-prived of reactants by a shutoff valve or pressure regulator which has failedclosed.3.3.2.1.10 FUEL-CELL AMMETER SWITCH.- A multiple-position rotary selector

switch is located Immediately below the fuel-cell ammeter. Thisselector switch allows current readout of each fuel-cell stack. The positionsare F.C. Stack 1A, IB, 1C, 2A, 2B, 2C respectively.


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REPORT.MODEL .

8638Gemini

3.3.2.23.3.2.2.1

SWITCHES - SILVER-ZINC BATTERY SUBSYSTEMMAIN BATTERY SWITCHES.- Five 2-position (ON-OFF) switchesare provided, one for each main battery. These switches permitconnection or removal of any main silver-zinc battery from the main bus.

3.3.2.2.2 SQUIB BATTERY SWITCHES.- Individual 2-position (ON-OFF) switchesare provided for control of the squib batteries. The commoncontrol bus is activated by throwing any of the squib battery switches. Squibbus 1 and 2 are activated by their respective control switches.REV

3.3.2.33.3.2.U

BUS TIE SWITCH.- Deleted.VOLTMETER SELECTOR SWITCH.- A multiple-position switch is providedfor applying the voltages obtained at various points in the elec-trical system circuits to the voltmeter (seeParagraph 3.3.1.4).

3.3.3 MISSION ELECTRICAL PROCEDURES.- During the course of a mission,periodic electrical system monitoring and power management isnecessary. In order that these functions may be punctually accomplished,routine checkout procedures are established. Tentative operational schedulesare categorized in accordance with mission phases and outlined in the followingparagraphs.3.3.3.1 PRELAUNCH.- During the prelaunch phase of the mission, while

ground power is utilized, bus voltages will be monitored andrelayed to the blockhouse on request. The fuel-cell battery system will beactivated and reactant quantity and pressure monitored. Switch positions fortransfer to spacecraft power will be selected and verified. Voltage and currentwill be monitored during transfer to spacecraft power, and launch-readiness ofthe electrical system verified.3.3.3.2 LAUNCH THROUGH INSERTION.- During this phase of the mission,voltage, current, hydrogen and oxygen quantity and pressure, andthe coolant pump warning lights will be monitored. In the event of unsatis-factory output from any portion of the electrical system, an attempt will bemade to restore the malfunctioning unit to normal operation. If this fails,the unit will be removed from the circuit, or a back-up method will be employed.In the event of a total fuel-cell subsystem failure, the fuel-cell subsystemand all non-critical equipment will be shut down and abort preparations made.


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6 June 1962 R E P O R T .MODEL .

8638Gemini

3.3.3.3 CATCH-UP THROUGH ORBIT.- After a successful insertion, the mainsilver-zinc batteries will be removed from the main bus. Per-iodic fuel-cell system purging will be accomplished as necessary to maintainoptimum output conditions. Battery conditions will be monitored periodically.Near the end of the final orbit, at approximately Tr-$ minutes, the silver-zinc batteries will be switched on and fuel-cell system power terminated.Voltage and current will be monitored during this transfer. Any power equip-ment malfunction during this phase of the mission will be treated in accordancewith the procedure outlined in Paragraph 33.32.3.3.3.1* RETROGRADE A N i J RE-ENTRY.- During the retrograde and re-entryphases of the mission, electrical conditions will be closelymonitored to insure that the system remains functionally secure throughoutincreased vibration and elevated heating periods. Override action will betaken if necessary in accordance with applicable procedures outlined in Para-graph 3.3.3.2.3.3.3.$ LANDING AND POST-LANDING.- During the landing phase of the mission,the electrical system outputs will be monitored and any necessaryoverride actions accomplished. Upon landing, all unnecessary electrically-powered equipment will be shut dovm to conserve power. Recovery equipmentwill be activated and system checks accomplished to determine availability ofnecessary post-landing power.

NEW


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MCDONNELL20 February 1962 ST. L O U I S , M I S S O U R I P A G E 1LR E V I S E D 6 June 1962 R E P O R T 8638R E V I S E D 15 November 1962 ,. fOmFIDEMTIAL * M O D E L Gemini

3.U CIRCUIT PROTECTION.- Electrical circuit protection is providedfor the purposes of fire protection and/or prevention of system Rwiring damage due to insulation failure. Protection is limited to areas inwhich input power is capable of producing current which is sufficient to damagethe wire. Circuit protection is accomplished through the use of sequenceswitching, fuses and/or circuit breakers, and reverse-current coils. Pyroswitches are provided for de-energizing interface circuits prior to severanceof spacecraft sections. Circuit breaker panels are provided which combine on-off switching provisions with circuit protection in some cases.


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D A T E 6 June 1962R E V I S E D 1*7 November 1962

ST. LOUIS, MISSOURI PAGE 15REVISED

REPORT.MODEL .

8638Gemini

3.53.5.1Report 8518.3.5.23.5.2.1

COMPONENTSREQUIREMENTS.- The general design requirements for electricalequipment used in the Gemini spacecraft are set forth in M.A.C,

BATTERIESFUEL-CELL BATTERY SUBSYSTEM.- Two fuel-cell battery sections arelocated in the equipment section of the adapter. Each sectionconsists of three stacks, contained within a pressurized sealed tank (seeFigure U). Each stack consists of a sufficient number of series-connectedcells to.maintain system voltage within established limits during all missionphases that require power produced by the fuel-cell batteries (see Figure 5)For normal operation, all stacks in a section are electrically connected inparallel to the main bus through individual ON-OFF switching provisions. Thefuel-cell battery system utilizes hydrogen and oxygen reactants stored in asupercritical cryogenic state (see Paragraph 3.5.2.1.1). Cooling of thefuel-cell battery system is provided by an active liquid cooling system. Thecooling system consists of two parallel loops, each having redundant pumps.Each loop circulates coolant through the cabin heat exchanger, suit heat ex-changer, equipment cold plates, and reactant and oxygen supply heat exchangersas well as the fuel-cell battery system. Each loop dissipates the accumulatedheat through an external radiator located on the surface of the adapter plusheat exchangers used to raise the temperature of the fuel-cell battery reactantsand environmental oxygen. For normal orbital periods one-pump, one-loopoperation is sufficient for cooling. During peak load period, both loops withtwo pumps, as required by the load, will be utilized. The potable water by-product of battery operation is transported to a central water reservoir locatedin the crew compartment of the re-entry module. The water reservoir provides avisual means of gauging the amount and condition of water contained plus a meansof positive removal of the water. Provisions are incorporated to allow purgingof inert gases from the fuel-cell battery as required. The total fuel-cellbattery system, including both fuel-cell sections, reactant supply tanks, andthe associated portion of the cooling system is installed in the adapter section.Any two of the six fuel-cell stacks are capable of supplying the basic electri-cal power requirement during orbital periods for both the missions specifiedin Paragraph 3 1> excluding rendezvous and docking phases.


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DA T ER EVI SEDR E V I S E D

20 February 19626 June 1962

15 November 1962

MCDONNELL.ST. LOUIS, MISSOURI P A G E


168638

GEMINI

FULL CELL B A T T E R Y S U B S Y S T E MJ8K SSttSI/awS:? ZK&f.l t t ?#.

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-

.--""'" ' ,$'& &\ \

0 * * * * * *//' r . i o w *u - - >x^-ift.} vJ.c. it"x ., r - .V J - A *---;".,3-|SSS

FIGURE 4


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C E N T E R P A U L


26/68

/ X

! o


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*fODO/VJViDATE 6 fyne3962 ST.LOUIS, MISSOURI PAGE lfi_REV1SED 11 July 1362 REPORT 8638R E V I S E D 1$November 1962 fPfirirTINT!I'l I M O D E L Gemini

3.5.2.1.1 REACTANT SUPPLY SUBSYSTEM.- The fuel-cell reactant supply subsystemis installed in the adapter section. This subsystem stores in afluid state, at supercritical pressures and cryogenic temperatures, the hydrogenand oxygen reactants which comprise the fuel and oxidizer utilized in theoperation of the fuel-cell battery subsystem. Reactant pressurization isaccomplished by individual heating elements contained within the H2 and 02tanks. These elements are segments of a closed-loop circuit which automaticallymaintains pressure within predetermined limits (SeeParagraph 3.32.1.7 forswitching provisions). The hydrogen and oxygen pass through heat exchangerswhich are components of the spacecraft cooling system. These heat exchangers,through which warm return coolant is circulating, warm the reactants to atemperature suitable for application to the fuel-cells. Sensors within thereactant supply subsystem provide discrete electrical signals for quantity andpressure in individual tanks and reactant temperature downstream from the heatexchanger (preheater). Materials used in the reactant supply subsystem whichare in direct contact with the oxidizer are not subject to pyrophoric reactions.3.5.2.2 SILVER-ZINC BATTERIES3.5.2.2.1 MAIN BATTERIES.- Five liO.p ampere-hour 16-cell mainbatteries of the silver-zinc type are provided. Open-circuitpotential of these batteries is 29.6volts maximum, with a plateau voltage ofapproximately 25.3 volts. Each battery is activated and sealed at sea levelpressure. Battery cases are vented to permit escape of gases and are providedwith suitable pressure relief valves. Cases are designed and constructed towithstand an interior-to-exterior differential pressure equal to 1.5 times themaximum relief pressure setting without permanent deformation or deformationadversely affecting battery operation. Venting provisions are designed to pre-clude electrolyte loss. The vent fitting and electrical connector are mountedas an integral part of the battery. Batteries are capable of operating in aweightless state, in any attitude except with cells upside down while underany acceleration force within the limits specified in Table II.3.5.2.2.2 SQUIB BATTERIES.- Three separate high-discharge-rate 16-cellsilver-zinc batteries are provided for supplying power to anisolated bus system. These batteries are isolated, both electrically andmechanically, from the other silver-zinc batteries and will maintain 16.0 voltsat the terminal under a 100-ampere load for one second.3.53 RELAYS.- The electrical relays used on the spacecraft are ofsufficiently rugged construction to withstand the mechanicalstresses, shocks, and vibrations characteristic of the planned mission. Theswitching mechanism is completely enclosed in a hermetically sealed cover, andis capable of operation in any position. Contact bounce is minimized to thebest value obtainable consistent with closure-time requirements. Relays shallbe of the immediate-closing, time-delay, and/or slow release type, dependenton the function performed in the individual circuit.


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D A T E o June lyo ST < L O U | S j M,SSOUR, PA GE 12_R E V I S E D 1^ November 1962 R E P O R T 8638

3.5.U SKETCHES3.5.1i.l TOGGLE SWITCHES.- Toggle switches used in the Gemini Spacecraftare sealed so as not to constitute a fire or explosion hazard,and are of construction sufficiently durable to withstand mechanical stressesimposed during the missions specified in Paragraph 3.1. Travel between extremepositions is between 26 and k6 degrees. Terminals are permanently and legiblymarked on the switch housing.3.5.b.2 ROTARY SWITCHES.- The rotary switches used in the Gemini spacecraftincorporate housings which completely enclose the switchingmechanism. The switches are sealed so as not to constitute a fire or explosionhazard. All terminals are solder-type, located on the back of the housing,A positive detent mechanism is provided, which minimizes the possibility ofthe movable contact coming to rest between contact positions. Contact arrange-ment is permanently marked on the switch houseing. Switches are constructedto withstand temperature, pressure, and acceleration variables encountered inthe course of the missions described in Paragraph. 3!3.5.U.3 LIMT SWITCHES.- Deleted. E3.5.U.ii TELELIGHTS.- The telelights used for functional indication on theGemini spacecraft incorporate opaque legends on translucentbackgrounds, with legends appearing dull black when the lamps are de-energized.The dual parallel-connected lamps are red, amber, or green as dictated bytelelight usage. Coloration is effected by the installation of filters on thelamps. Test terminals are provided, with an internal diode arrangement toprevent1 energizing the input circuit when test voltage is applied.3.5.it.U.I TELELIGHT/SWITCHES.- The telelight/switches incorporate thefeatures described above, and in addition provide an integralswitching capability. The positive-break switch is housed in a hermetically-sealed case and actuated when the telelight display is pressed. Switches areof the momentary-contact type, which automatically return to normal positionupon release, and alternate-contact type, in which each successive actuationresults in repositioning to the alternate set of contacts.3.5.5 EXTERNAL POWER CONNECTIONS.- Prior to launch, external power issupplied to the re-entry module to prevent undue depletion of thespacecraft power supply. This is accomplished by means of an electrical dis-connect assembly. The adapter and re-entry module also contain electricalconnections for monitoring spacecraft parameters during prelaunch checkout.Each disconnect assembly consists of a plug and receptacle assembly and acoupling device for securely retaining these components in the engaged position.The disconnect assembly is capable of positive disengagement/separation bothelectrically by a solenoid device and manually by a lanyard-initiated mechanism.The re-entry module disconnect is normally released at approximately T-30seconds. The adapter module disconnect is normally released at T+3 to T+5seconds, with the backup device actuated by movement of the launch vehicle.

RV

REV


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D A T E 6 June 1962 ST. LOUIS, MISSOURI P A G E 20REVISEDREVISED

November 1962 REPORT.MODEL .

86^8Gemini

3.5.6 CONNECTORS.- Miniature bayonet (Bendix Pygmy) type connectors areused throughout the electrical subsystem, except at external powerand interface connections. See Paragraph 3*5.5 for external power connectors

and M.A.C./Martin Report 86l2/MB-10li8 for interface connector specifications.3.5.7 CIRCUIT BREAKERS.- The circuit breaker/switches used in theGemini Spacecraft are of the magnetic type. Choice of this typeis predicated on its relative insensitivity to thermal changes, which enhancesstability characteristics through temperature variations encountered in orbitalmissions. Hermetically-sealed integral switches are incorporated which providecapability of complete deactivation of selected circuits.3.6 FABRICATION.- The methods and procedures utilized in the fabri-

cation of the electrical system components and subassemblies arein accordance with applicable M.A.C. Process Specifications as listed in Para-graph 2.2.2, Government specifications insofar as practicable, and establishedspacecraft standards. Only materials of highest quality are used. Temperature-resistand and insulating materials are used wherever necessary to insure ful-fillment of the electrical system's function in contributing to completion of asuccessful mission. Miniaturization and weight reduction are emphasized insystem design. Wherever possible, existing components and items of equipmentare used to reduce the lead time necessary for development of the system. De-velopment-type items are obtained through liaison with, procurement from, and/orassistance of, reputable manufacturers. Spare pins, with short wire segmentsattached, are provided in various connectors throughout the electrical systemto provide for growth, routing changes and/or special circuits. Test connectorsare incorporated at various points to facilitate connection of aerospace groundequipment during spacecraft systems testing and prelaunch checkouto

REV


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MODOJVAURl PAGE 86 8

D A T E 6 June 1962 - ST> LOU1S> M|SSOUR, PAGE 21_REVISED REPORTREVISED MJbOvUA LMHM MODEL

3-7 INSTALLATION.- Installation of system components and equipmentis in accordance with applicable M.A.C. Process Specifications aslisted in Paragraph 2.2.2, contractor and vendor engineering drawings, andGovernment standards insofar as practicable. Location of items is based onfunction performed, associated equipment, communications interference, environ-mental adaptability, heat generation, space limitations, accessibility, andcrew station compatibility. Possibility of shorting due to moisture, corrosiveatmosphere, floating debris, and/or detached equipment is minimized by elimina-tion of exposed powered electrical connections throughout the system. Thisis accomplished by insulating and/or sealing terminals, connections, splices,bare conductor sections, lugs, studs and solder pots with potting, coating, orsealing compounds, paint, tubing, sleeving, plastic foam and/or mechanicalcovers. Wherever feasible, non-captive wire bundle installation is utilized,i.e., removal of wire bundles may be accomplished without removing or separatingsoldered or other permanent-type connections.


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DATE * June 1962 ' ST. LOU.S. M . S S O U R , P A G E 22.8638REVISED R E P OR T

k. ENVIRONMENTAL CONDITIONSU.1 GENERAL.- This system and its components vill "be capable of .acceptable performance when subjected to the environmentalconditions as specified in Tables I, II or III, as applicable, except asotherwise specified in individual paragraphs herein.5 QUALITY ASSURANCE51 GENERAL.- System components and materials are subjected to bothacceptance tests and design approval tests. Acceptance testsinclude, but are not limited to, those tests performed to insure that thematerials, workmanship, and performance of components are not substandardand that the components have been manufactured to approved drawings andspecifications. Design approval tests are defined as those tests conductedon preproduction and/or production equipment to determine that the designof the equipment complies with all requirements therefor. Production-linesurveillance is maintained during fabrication, assembly and installation ofall components and subassemblies to insure that all methods, procedures andambient conditions are maintained in accordance with existing specificationsand established spacecraft standards. Quality assurance provisions are inaccordance with M.A.C. Report 8580-7, "Quality Assurance Provisions (Plan)for Project Gemini Space System".6. RELIABILITY6.1 GENERAL.- An active reliability program is established andconducted throughout the design, development, fabrication, andinstallation of the Gemini electrical system. Adequate testing, both in-plantand/or vendor/subcontractor-accomplished, is performed to insure a reliabilityfactor which does not deteriorate the overall probability of a successfulmission. Mission safety is emphasized in the design approach, back-up andredundant systems are utilized where necessary, and every effort is made tominimize system complexity. Reliability provisions are in accordance withM.A.C. Report 8580-3, "Project Gemini Reliability Plan."7. DATA REQUIREMENTS71 GENERAL.- Applicable design information and performance reportsare submitted in accordance with the requirements specifiedin M.A.C. Report 8580-8, "Model 133P Gemini Program Documentation Plan."


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6 j 1Q62 flf OOOHJ 2 ST. LOUIS. MISSOURI PAGE 23

REVISED ; REPORT 8638R E V I S E D &$HP$&9t4*M M O D E L Gemini

8. TESTING8.1 GENERAL.- In addition to the acceptance, design approval, andreliability tests described in Paragraph 5.1 and 6.1, the elec-trical system will undergo numerous other tests, both as an individual systemand as an integrated portion of the spacecraft system as a whole. Details ofthe various tests are itemized in M.A.C. Report 8580-5. Basic test groups aredescribed in the following paragraphs. Throughout the testing program, systemperformance and interaction with other systems will be monitored, recorded,and evaluated. Design, fabrication, procedural, and/or installation correctionsand improvements will be implemented where necessary as a result of theseevaluations.8.1.1 BATTERY TESTS.- A prototype fuel-cell battery system (lessreactant tank's) will be assembled and tested for fuel flow, purgerequirements, voltage regulation, and susceptibility to voltage fluctuations.This prototype system will be upgraded to a production system as parts becomeavailable, and installed in the compatibility test unit, where it will be Nemployed to power the spacecraft systems in a simulated mission evaluating the Ebatteries under actual spacecraft systems loads and transients. Switchover Wfrom fuel-cell to silver-zinc batteries will be accomplished during this oper-ation. The reactant storage system will be separately tested under criticalfuel flow.

The fuel-cell battery/reactant storage system will be installedin an altitude chamber and tested under simulated orbital load (which will bedetermined from tests performed with the spacecraft systems of the compatibilitytest unit) until the reactant supply is totally depleted, to determine theampere-hour capacity for a single fueling.8.1.2 COOLING SYSTEM TESTS.- The common cooling system which servesthe fuel-cell power supply, equipment cold plates, and environ-mental system heat exchangers will be tested in production prototype form ina simple mockup which simulates component spacing and geometric relationship.Heating elements will simulate equipment heat loads and the production proto-type power supply or heating elements will provide the fuel-cell battery heatload. The tests will evaluate flow rates, pressure drops, and heat transfercharacteristics of the radiator and cold plate/circulation systems, bothseparately and combined, under ambient and simulated altitude conditions. Ihecomplete production prototype system will be installed and tested in thecompatibility test unit and with the production prototype fuel-cell system(see Paragraph 8.1.1). The complete spacecraft cooling system will be testedunder simulated mission environmental conditions during the course of thesimulated orbital mission (see Paragraph 8.1.1;).


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.T . . MOOOAfJVJD A T E b. June 19&2 ST> LQU|S> M1SSOUR1 PAGE8638REVISED REPORT

8.1.3 SPACECRAFT SYSTEMS TESTS (SST).- The production electrical systemor sections thereof, as required, vdll be utilized in conductingoverall spacecraft systems tests, including individual system operational tests,integrated system operational tests, and simulated mission tests. These testswill be performed on production spacecraft systems which are in sufficientstages of completion to permit accurate evaluation of data relative to deliveredspacecraft configuration. E

W8.1.U SIMULATED ORBITAL MISSION.- A complete electrical system will beinstalled in a spacecraft which will be used in a completesimulated mission, run at reduced.pressures to simulate altitude, elevatedtemperatures to simulate launch and re-entry heat loads, and reduced temperaturest o simulate orbital conditions. Sequential operation of the electrical systemwill be accomplished, and all performances under extreme environmental conditionswill be monitored, recorded and evaluated, and any unsatisfactory conditionscorrected.


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.D A T E _6 June 1962 _ $TtLOU1S> M|SSOUR, PAGE _ 25REVISED _ REPORT _ 86"38REVISED_ COJMPTOHMODEL_Gemn

9. DEFINITIONS9.1 GENERAL.- The terms listed below are used, in this specificationand are defined as follows:

Spacecraft - The adapter module and re-entry module. Thisterm applies either with or without the launch vehicle mating section of the adapter module.

Adapter - That portion of the spacecraft used primarilyto support the re-entry module on the launch vehicle.Re-entry Module - That portion of the spacecraft whichhouses the crew. The part of the spacecraft which re-enters \the earth's atmosphere.Tr-AT - An arbitrary time prior to retrograde firing,just prior to equipment-section jettison, at which timere-entry preparations and sequencing are normally initiated.Equipment Section - The central portion of the adapterwhich houses systems used during the orbitalphase of the mission; jettisoned prior to retrogradefiring.


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D A T ERE VI SE DREVISED

MCDONNELLST. LOUIS. MISSOURI

ENVIRONMENTAL CONDITIONSEquipmentLife

EnvironmentAmbient TemperatureAnbient PressureTemperature-PressureRelative HumidityRainSalt Sea AtmosphereSand and DustFungus (2)O a g r g e n AtmosphereShock (3) (U) (5)

Acceleration (3) (U) (6)

Vibration Equipment Category A (Figure 6) (7)Equipment Category B (Figure ?) (8)Random (Figure 8)Acoustic NoiseRadio InterferenceAtmosphere

Pre2023.5N . A

MILMILMIUM I L100*N.A

N.A

ProProtProProM I LI.A.

NOTESt1. N.A. Not applicable.


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PAGE _REPORT.MODEL .

268636L

TABLE I IN THE RE-ENTRY MODULE IN THE

Operation160F15.5 psia

Procedure IIProcedure IProcedure IProcedure I

at 15.5 psia

Launch0F to 160F15.5 to 10-12 psia16C?F at 15.5 to lO'12psia15* to 100*N . A ,N . A ,N . A ,N . A .lOCtf Og at 6,0 psiaN . A ,

Longitudinal Space-craft Axis: 1. gto 7.25 g linearlywitn time over 326 sec.Curve ICurve ICurve 1135 db Over-allSee Figure 9

Or0F to 16.0 to 1C160F at5C# to 8N . A .N . A .N . A .N . A .100$ 02 N . A . (1)

Og

Curve IICurve IICurve IIN.A.


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CABIN

Re-Entry Post-Landing

Dsiapsia

psia

200F10-12 to 15.5 p200F at1C'12 to 15.5psia over 10 HLn,15* to 10056N.A.N.A.N.A.N.A. 03 at 6.0 psiaSee Figure 13

Longitudinal SpacecraftAxis: I5g, 30second duration.Lateral Spacecraft Axis;U.5 g, 30 sec, durationCurve IIICurve IIICurve IIISane as for LaunchMIL-I-26600N.A.

-15 to 160F15.5 psiaN.A.15* to 100$N.A.N.A.MIL-E-5272CProcedure IN.A.15 g's in anydirection11 millisecondduration1 g R

EV

N.A.N.A.N.A.N.A.KIL-I-266CON.A.


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DATE Jam 1O9REVISEDREVISED

MCDONNELLST. LOUIS. MISSOURI

ENVIRONMENTAL CONDITIONS

Environment Prel&unch OperationAmoient TemperatureAmbient PressureTemperature-PressureRelative HumidityRainSalt Sea AtmosphereSand and DustFungus (3)Shock (U) (5) (6)

Acceleration (U) (5) (8)

VibrationCategory A (Figure 6) (9)Category B (Figure 7) (10)Random (Figure 8)Acoustic Noise (11)Radio InterferenceExplosive Atmosphere

20F to 160F1$.5 psiaN.A. (1)15* to 100*MIL-E-5272C, Procedure IIMIL-E-5P72C, Procedure IM1L-E-5272C, Procedure IMIL-E-5272C, Procedure IN.A. (1)

N.A. (1)

protectedProtectedProtectedProtectedMJL-I-26600B.A.


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.9

TABLE IIELECTRICAL EQUIPMENT LOCATED IN THE RE-ENTRY M O D U L E EXTERNAL TO

Launchto 16QOFto 10-12 pgiaat 10-12 psia

* to 100*( 1 )(1)(1) |( 1 )(1)

spacecraft axis:Ig to 7.25g, linearly withtime over 326 seconds.

IIIS db Over-all, See Figure 10

( 1 )

Orbit-60F to 160F160F at 10-12 psia160F at 10-12 psiaN . A . (1 )N . A . (1 )N.A. (1)N . A . (1)N . A . (1)N . A . (1)

Og

Jurve IICurve IIJurve IIN . A . (1)KIL-I-26600N . A . (1)

200F10-1 2 to 1200F at 1C10 minut15* to looN . A . (1)N . A . (1)N . A . (1)N . A . (1)Area I -3Ccraft, 3spacecraArea II Longitudin30 secomspacecrdurationCurve IIICurve IIIC u rv e IIISame as foMIL-I-2660CN.A. (1)


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-PRESSURIZED CABIN

Post-Landing

psiato 15.5 psia over

11 ms duration (7)Figure 13 (7)I5g,

30 seconds

-15F to 160F15.5 psiaN.A. (1)

t oMIL-E-5272C, U-hr.Immersion (2)MIL--5272, Procedure IMIL-E-5272C, Procedure IMIL-E-5272C, Procedure II5g's in any direction 11 milli-second duration.

lg

N.A. (1)N.A. (1)N.A. (1)N.A. (1)MIL-I-26600

RV


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DATEREVISEDREVISED

MOOOJVJVJST. LOUIS. MISSOURI

ENVIRQEquipmentLife^ PhaseEnvironnent P

A m b i e n t TemperatureAmbient PressureTemperature-PressureRelative HumidityRainSalt Sea AtmosphereSand and DustFungus (2)Explosive AtmosphereShock (3) (U) (5)Acceleration (3) (U) (6)

VibrationCategory A (Figure 6) (7)Category B (Figure 7) (8)Random (Figure 8 )Acoustic Noise (9)Radio Interference

20F to16015.5 psiaN.A. (1)1# to 100MIL-E-5272CMIL-S-5272CMIL-E-5272CHydrogen AtN.A.N.A.

ProtectedProtectedProtectedProtectedMH-T-26600

NOTESt


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.2JL

8638Gemini

TABLE IIICONDITIONS FOR ELECTRICAL EQUIPMENT LOCATED IH THE ADAPTER

Launch

IIIII(10)

-60F to 160F15.5 to 10"12 psia160F et 15.5 to 10'12 psia15* to 100*N.A.N.A.N . A .N . A .N. A.N . A .Longitudinal Spacecraft Axisl.g to 7.25g linearly with tover 326 sec.Curve ICurve IC U T T C I155 db Over-allSee Figure 11MIUT-26600


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Orbit

-60F to 160F10"12 psia160F at 10"12 psiaN.A.N.A.N.A.N.A.N.A.N.A.N.A.Og

Curve IICurve IICurve IIN.A.MIL-I-26600

hV


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D A T E 6 June 1962R E V I S E DREVISED

MODOJVJVfJLJLST. LOUIS, MISSOURI P A G E

REPORT.MODEL .

298638

Gemini

GEMINI SPACECRAFT VIBRATION SPECTRAEQUIPMENT C A T E G O R Y A

.00001F R E Q U E N C Y (CPS) Figure 6


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DATE 6 June 1962 ST. LOUIS, MISSOURI P A G EREVISEDREVISED

REPORT.MODEL .

308638

Gemini

GEMINI SPACECRAFT VIBRATION SPECTRAEQUIPMENT CATEGORY B

i.o

.10

.01u

"'

.0001

.00001

TO 5 CPS

CURVE. CURVE II

CURVE III

\\V

10 50 100 500FREQUENCY (CPS)

1000 5000

Figure 7


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D ATE 6 June 1969 MCDONNELLr. LOUIS, MISSOURIREVISEDREVISED MODEL

318638

ZO|ttt


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D A T E 6 June 1962_ ST. LOUIS, MISSOURI PAGE 32REVISEDREVISED

REPORT.MODEL .

8638Gemini

(/zUJo2UluOu


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D A T E A Jnne 1962

REVISED

MGDO/VJVjEST. LOUIS, MISSOURI P A G E 8638

MODEL

LLJ_Ji> U 2O


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D A T E 6 June 1962RE V I S E DREVISED

S T. LOUIS, MISSOURI P A G E 3ilR E P O R T ,MODEL .

8638Gemini

ttC5

Z iSui a:

/> 1--

t2idu

jWD/S3NAQ ZOOO1 :3M SO - 13A31 3&nSS3Ud ONOOS

Figure 11O O l i r i D D N T L I L >


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6 June 196 MGOOJVJVST. LOUIS, MISSOURIR E V I S E DR E V I S E D

REPORT.MODEL .

UJi3U zouiZio so-O-

T

1 1 I 1 1 1 1 1 I I 1 1 I 1 1 I 1 I 1 1 1 1 1 I I I I 1 1 I 1 1 I 1 1 I 1 1 1 1 1 1 1 I I

W D / S 3 N A O Z O O O - =38 90 - T 3 A 3 T 3 d f \ S S 3 8 d

So0

8

o oS3

a.o />a

88 Irn o ...

aUJ

Figure 12


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D A TE 6 June 1962 ST. LOUIS, MISSOURI P A G E 36REVISEDR EVI SED


8638Gemini

. .< 86

o

Figure 13


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DA T E 6 June 1962 MODOJVAUEST. LOUIS, MISSOURIREVISED 11 .Tnly 1962REVISED

P A G E 378638

MODEL Gemini

APPENDIX IE Q U I P M E N T LIST .

Qty. Nomenclature1 Battery Subsystem - Fuel Cellk Battery - Silver Zinc3 Battery - Silver Zinc* Relay - Control* Relay - Time Delay# Switch - Toggle* Switch - Rotary* Switch - Limit1 Umbilical - Ground1 Umbilical - Ground* Umbilical - Staging* Telelight

= * Telelight/Switch* Telelight/Switch

t* Relay - Voltage Sensing* Fuse/Fuse Resistor* Circuit Breaker/Switch* Shunt

* These quantities will be inserted as f i rm

M . A . C . No.52-79700-352-79702-352-79702-552-7970352-7970152-7970552-79706 I52-79707 *52-79708-352-79708-552-7970952-7971052-7971052-7971052-7971552-7972052-79721

requirements are determined.


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DATE 6 June 1962 MODOJVJVfiST. LOUIS. MISSOURI P A G E 38REVISEDR EVI SED


8638Gemini

Qty.

#

APPENDIX I (Continued)EQUIPMENT LIST

NomenclatureSystem - Reactant SupplyExtended MissionSystem - Reactant SupplyShort Mission

M.A.C. No.52-83701-1

52-83701-3NEW

# One of these systems is used for each spacecraft, dependent on themission for which it is utilized.


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DATE 6 June 1962 MODOJVJVCST. LOUIS. MISSOURI P A G E 39REVISEDREVISED

11 July 1962 R E P O R T .MODEL .

8638Gemini

APPENDIX IICHANGE LIST

Following is.an itemized list of changes incorporated in McDonnellReport 8638 as a result of successive revisions. In instances where con-flict or confusion may result, the authority for these changes is parenthe-tically specified; where the reason for change is self-explanatory (e.g.,rearrangement of format, addition of reference material, rewording forclarity or additional information, etc.), no authorization is listed.

Section 1

References:ParagraphTitle Page

Changes Incorporated in 11 July 1962 RevisionA. NASA-M.A.C. Houston Co-ordination Conference - 25-27 April 1962

Changes

Page ii and iii

Page ivPage vPage vi1.12.2.1

2.2.2

3.1

Changed "Subsystem Specification" to "SystemSpecification". (Reference A)Added pages i through vi; revised Table ofContents to reflect present format; addedAppendices I and II.

*Added Figures 2 through 13.Added Pages 15 through 1*1.New page.Added "Testing".Rearranged - added McDonnell Reports 8392,8580-5, and 8610. Added McDonnell Drawings52-79706, 52-79709, 52-79710, 52-797H,52-79715, 52-79720, and 52-83701.Removed P.S. 17305 and P.S. 17 10.3. (Notapplicable to electrical system.)Added..."with .a minimum of interferencebetween the various subsystems and devices."Added..."2-Wire Grounded System". AddedParagraphs outlining AC power and intercon-necting wiring philosophies. (Reference A)

NEW


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D A T E 6 June 1962R E V I S E D 11R E V I S E D

ST. LOUIS, MISSOURI P A G EREPORT.M O D EL .

8638Gemini

Paragraph3.2

3-3

3 - 5

.3.5.13.5.2.13-5.2.2

3-5.33.5. .13.5. .23-5.^.3

APPENDIX II (Continued)CHANGE LIST

ChangesParagraph title was: "Fabrication", now:"Power Distribution"; content changedaccordingly.Paragraph title was: "Installation"; now:"System Management"; content changed accord-ingly. This paragraph and all subparagraphsthereto are new paragraphs. (Reference A)Paragraph title was "Power Distribution";now: "Circuit Protection"; content changedaccordingly. .Expanded paragraph to includecircuit protection philosophy and methods.(Reference A)Paragraph title was: "Monitoring"; now:"Components"; content changed accordingly.With the exceptions noted in the followingparagraphs, this section is the same asSection 3.7 in the basic issue.

4New Paragraph.Reworded to update terminology.Reworded battery description; revised ratingsto reflect later information.Added voltage-sensing relays.Added toggle travel and terminal marking.New Paragraph.Removed toxic fumes/noxious odors statement(covered by McDonnell Report 8518; seeparagraph 35.1 - Requirements.New Paragraph.


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DA T E 6 June 1962REVISEDREVISED

ST. LOUIS, MISSOURI P A G ER E P O R T .MODEL .

8638Gemini

Paragraph3.5-5

3-5.63.6

3.7

U.lSection 8

Section 9

APPENDIX II (Continued)CHANGE LIST

ChangesReworded; expanded to include approximaterelease times.New Paragraph.Paragraph title was: "Circuit Protection";now: "Fabrication"; content changed accord-ingly.Paragraph title was: "Components"; now:"Installation"; content changed accordingly.Expanded to add information.Added... "as applicable".... after tables.Section title was; "Definitions"; now:"Testing". Content of entire section new.(Reference A)New Section No. - Content same as Section8 in basic isue.Revised Tables I, H> and III to reflectlatest information; added Figures 6through 13 as an adjunct to these Tables.Added Appendices I and II.


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DATE 1S NovemberREVISEDREVISED .

ST. LOUIS, MISSOURI PAGEREPORT.

MODEL .

8638

Section 2Changes Incorporated in 1$ November 1962Revision

ParagraphPage ii and iii

Page vi

2.32.3-12.3-22.3.32.k3 - 13.1.13-1.23-1.33.23-2.13.2.23.3.1.23.3.1.33 - 3 . l . 1*3-3.1.73-3.1.8

3-3.1.93-3.2.1.33.3.2.1.if3-3.2.1.63.3.2.1.73.3.2.1.8

ChangesRevised Table of Contents to reflect newparagraphs.Incorporated 15 November 1962 revision inIndex of RevisionsRearranged and renumbered from 2.2.1and 2.2.2

Renumbered from 2.3Divided 3-1 into separate paragraphs andchanged paragraph reference from 3'5-1 to 3 - 5-2

Divided 32 into separate paragraphs andupdated content

Deleted as requirement no longer existsUpdated content

Changed paragraph reference from 3-32 to3-3.2.1.6Added in accordance with updated designUpdated contentDeleted as requirement no longer existsUpdated content

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DATE 15 November 1962 ST. LOUIS. MISSOURIREVISEDREVISED

PAGE ->REPORT 8638MODEL Gemini

Paragraph3.3.2.1.93-3.2.1.103-3.2.2.13.3.2.2.23-5.2.13.5.2.2.2

3.5.33.5-53.5.6

Fig. 1Fig. 2Fig. 3Fig. k

Addendum A

Section 2 - (Continued)Changes

Added in accordance vith updated design

Updated content

Updated content and renumbered fromparagraph 3-52.2Updated content

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Added to reflect Spacecraft Noelectrical system.


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D A T E 15 November 1962REVISEDREVISED :

ST. LOUIS, MISSOURI PAGE AlREPORT.MODEL .

~ Add- AGemini

ADDENDUM AELECTRICAL SYSTEM - SPACECRAFT NO. 1

NW


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Al11.1A2A2.1A2.2A2.2.1A2.2.2A2.3A3A3.1A3. 1.1A3.1.2A3. 1.3A3. 2A3. 2.1A3. 2. 2A3 .3A3. 3.1A3. 3. 1.1A3.3 - 1.2A3. 3. 2A3.3.2.1A3. 3. 2. 2A3.3.2.313.1*A3. 513.5-1A3.5.2A3. 5. 2.1A3.5.2.2A3. $.313.5.1*13.5.1*.!A3. 5.1*. 21 3 . 5 - 1 * 313.5.1*.!*A3.5.5A3.5.6A3. 5. 7A3. 6A3. 7

TABLE OF CONTENTSTitle

TITLE PAGEGENERALScopeAPPLICABLE DOCUMENTSGeneralDocumentsMcDonnell DocumentsProcess SpecificationsSupersedenceDESCRIPTIONGeneralPower SourceAC PowerWiring and ConnectorsPower DistributionMain Bus SubsystemIsolated Bus SubsystemSystem MonitoringVisual IndicatorsAmmeterVoltmeterSwitchesMain Battery Switch.Squib Battery SwitchVoltmeter Selector SwitchCircuit ProtectionComponentsBatteriesBatteriesMain Silver-Zinc BatteriesSquib BatteriesRelaysSwitchesToggle SwitchesRotary SwitchesLimit SwitchesTelelight/SwitchesExternal Power ConnectionsConnectorsCircuit BreakersFabricationInstallation

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11*Al*Al*HiAl*11*A5A5A 6A 6A6A6A6A6A6A6A6A6A?A7A 7A7A?A?A?A?A?A 7A7A8A8A8A8A8A8A8A9A9A9A9A9


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8636 - Add.Aflonri r> AGemini

AlAl.l

GENERA!SCOPE.- Addendum A, which is the same as the basic specificationexcept as specified, defines the technical requirements for design,development and testing of the Electrical System for Spacecraft No. 1 of ProjectGemini to be constructed by McDonnell Aircraft Corporation under Contract NAS9-170 for National Aeronautics and Space Administration (NASA). This documentis not to be construed as, or used for ai inspection guide.

A2A2.1A2.2A2.2.1

APPLICABLE DOCUMENTSGENERAL.- Sane as basic specification.DOCUMENTS.- Same as basic specification.MCDONNELL DOCUMENTS.- The following McDonnell documents areapplicable to this addendum insofar as their content concerns thesystem described herein:Report 8357

Report 8518

Report 8080-3Report 8580-5Report 8580-7

Report 8580-8

Report 8610

Report 8611

Drawing 52-79702Drawing 52-79703Drawing 52-79701*

"Gemini Spacecraft ApplicableDocuments""Electrical and E3e ctronic EquipmentDesign Requirements for Model 133P""Project Gemini Reliability Plan""Project Gemini Test Plai""Quality Assurance Provisions (Plan)for Project Gemini Space System""Model 133P Gemini ProgramDocumentation Plan""Gemini Spacecraft EnvironmentalCriteria Specification"Gemini Spacecraft PerformanceSpe ci f ica ti on"Battery, StorageRelays, Control and PowerRelays, Electronic Time Delay


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R E P O R T 86^8 - A d d . AM O D E L Gemini

A2.2.1

A2.2.2A2.3

MCDONNELL DOCUMENTS.- (Continued)Drawing 52-7970$Drawing 52-79707Drawing 52-79708Drawing 52-79709Drawing 52-79710Drawing 52-79711Drawing 52-79715Drawing 52-79720Drawing 52-79721Drawing 52-90000

Switches, ControlSwitches, limitGround UmbilicalStaging UmbilicalTelelights and Telelight SwitchesFloodlightRelay, Voltage SensingFuse/Fuse ResistorCircuit BreakerFinish Specification

NE"

PROCESS SPECIFICATIONS.- Same as basic specification.SUPERSEDENCE.- Same as basic specification.


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8638 - Add. AGemini

A3A3.1DESCRIPTIONGENERAL.- The function of the Gemini spacecraft electrical systemis to supply and distribute electrical power of the proper voltageto all spacecraft devices requiring electrical power for operation, with aminimum of interference between the various subsystems and devices.

A3.1.1 POWER SOURCE.- The electrical system is a 2-wire, grounded systemutilizing silver-zinc batteries as sources of DC power. Thebatteries shall be located in the re-entry module of the spacecraft, outsidethe pressurized area.A3.1.2 AC POWER.- Devices utilizing AC power are supplied by self-contained inverters within the individual subsystems. These in-verters are not considered a part of the electrical system, but are treated asinherent elements of the subsystem which they serve.A3.1.3 WIRING AND CONNECTORS.- For purposes of this specification, thewiring and connectors which interconnect components of systemsother than the electrical system are considered a part of the electrical systemand shall meet all applicable requirements specified herein.A3.2 POWER DISTRIBUTION.- DC power distribution to the utilizing systemsis accomplished through a main bus subsystem and an isolated bussubsystem, each of which is independent of the other.

NE

A3.2.1

A3.2.2

MAIN BUS SUBSYSTEM.- The main bus subsystem consists of a mainbus energized by four silver-zinc batteries.ISOLATED BUS SUBSYSTEM.- The isolated bus subsystem consists oftwo squib buses and a control bus energized by three isolatedsilver-zinc batteries throughout the mission. The functions of this subsystemare to supply power to the pyrotechnic ignition devices through redundant,isolated, parallel circuits from power source to igniter and to supply powerto solenoids and control relays. The control bus is energized by both squibbuses and a separate battery, but isolated from each by diodes. In the interestof launch site, and in-flight safety, provisions are incorporated for disarmingthe pyrotechnic circuits. Redundant pyrotechnic control is provided.

A3.3 SYSTEM MONITORING.- Provisions are incorporated on the instrumentpanel for monitoring the power sources. Monitoring of the systemwill be followed throughout the mission by camera and recorders.A3.3.1 VISUAL INDICATORS.- Visual indications of electrical system events,conditions, outputs and malfunctions are provided, consisting ofwarning and indicator lights and single and multiple reading meters.


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A3.3.1.1 AMMETER.- A dual-reading ammeter is installed on the instrumentpanel for monitoring current flow of the main batteries. Batteries1 and 2 are indicated on one scale with 3 and k on the other scale.

A3.3.1.2 VOLTMETER.- A voltmeter is provided on the instrument panel formonitoring individual voltages at the main bus, squib buses andthe control bus.A3.3.2A3.3.2.1

SWITCHESMAIN BATTERY SWITCH.- Four 2-position (ON-QFF) switches are pro-vided, one for each main battery. These switches permit connectionto or removal of any main silver-zinc battery from the main bus circuit.

A3.3.2.2 SQUIB BATTERY SWITCH.- Individual 2-position (ON-OFF) switches areprovided for control of the squib batteries. The common controlbus is activated by throwing any one of the squib battery switches. Squibbuses 1 and 2 are activated by their respective control switches.A3.3.2.3 VOLTMETER SELECTOR SfclTCH.- A multiple-position switch is pro-vided for applying the voltages obtained at various points in theelectrical system circuits to the voltmeter.A3.il CIRCUIT PROTECTION.- Electrical circuit protection is provided forthe purpose of fire protection and/or prevention of system wiringdamage due to insulation failure. Protection is limited to areas in whichinput power is capable of producing current which is sufficient to damage thewire. Circuit protection is accomplished through the use of sequence switching,fuses and/or circuit breakers, and reverse-current coils.A3.5A3.5.1Report 8518.A3.5.2

COMPONENTSREQUIREMENTS.- The general design requirements for electricalequipment used in the Gemini spacecraft are set forth in McDonnell

BATTERIESA3.5.2.1 MAIN SILVER-ZINC BATTERIES.- Four UO.O ampere-hour 16 cellmain batteries of the silver-zinc type are provided. Open circuitpotential of these batteries is 29.6 volts maximum, with a plateau voltage ofapproximately 25.3 volts. Each battery is activated and sealed at sea levelpressure. Battery cases are vented to permit escape of gases and are providedwith suitable pressure relief valves. Cases are designed and constructed towithstand an interior-to-exterior differential pressure equal to 1.5 times themaximum relief pressure setting without permanent deformation or deformationadversely affecting battery operation. Venting provisions are designed to pre-clude electrolyte loss. The vent fitting and electrical connector are mounted

NEV

MA C 23 1CM (REV 14 JUN 62!


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BlfTIALREPORT 8638 - Add. AM O D E L Gemini

A3.5.2.1 MAIN SILVER-ZINC BATTERIES.- (Continued)as an integral part of the battery. Batteries are capable of operating in aweightless state, in any attitude except with cells upside down while under anyacceleration force within the limits specified in Table II.A3.$.2.2 SQUIB BATTERIES.- Three separate high-discharge rate 16-cellsilver-zinc batteries are provided for supplying power to theisolated bus system. These batteries are isolated, both electrically andmechanically, from the main silver-zinc batteries and will maintain 16.0 voltsat the terminal under a 100-ampere load for one second.A3.5.3 RELAYS.- The electrical relays used on the spacecraft are ofsufficiently rugged construction to withstand the mechanicalstresses, shocks, and vibrations characteristic of the planned mission. Theswitching mechanism is completely enclosed in a hermetically sealed cover, andis capable of operation in any position. Contact bounce is minimized to thebest value obtainable consistent with closure-time requirements. Relays shallbe of the immediate-closing, time-delay or slow dropout type, dependent on thefunction performed in the individual circuit.A3.5.U SWITCHESA3.5.1ul TOGGLE SWITCHES.- Toggle switches used in the spacecraftare sealed so as not to constitute a fire or explosion hazard, andare constructed to withstand mechanical stresses imposed during the mission.A3.5.U.2 ROTARY SWITCHES.- The rotary switches used in the spacecraft in-corporate housings which completely enclose the switchingmechanism. The switches are sealed so as not to constitute a fire or explosionhazard. All terminals are solder-type, located on the back of the housing. Apositive detent mechanism is provided, and contact arrangement is permanentlymarked on the switch housing. Switches are constructed to withstand temperature,pressure, and acceleration variables encountered during the course of themission.A3 !*.3 LIMIT SWITCHES.- Limit switches used in the spacecraft are sealed

to eliminate fire or explosion hazards in an oxygen atmosphere asencountered in the pressurized cabin of the spacecraft. The limit switches arecapable of functioning under pressure values and extreme temperature variationsas dictated by the mission.A3.5.U.1* TELELIGHT/SWITCHES.- The telelight/switches used for functionalindication on the spacecraft incorporate opaque legends ontranslucent backgrounds and an integral switching capability. The dualparallel-connected lamps are red,amber, or green as dictated by telelightusage. Coloration is effected by the installation of filters on the lamps.

NEW

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A3.5.U.U TELELIGHT/SWTTCHES.- (Continued)Test terminals are provided, with an internal diode arrangement to preventenergizing the input circuit when test voltage is applied. The positive-breakswitch is housed in a hermetically sealed case and actuated when the telelightdisplay is pressed. Switches are of the momentary-contact type, which auto-matically return to normal position upon release and alternate-contact type inwhich each successive actuation results in repositioning to the alternate setof contacts.A3.5.5 EXTERNAL POwER CONNECTIONS.- Prior to launch,, external power issupplied to the adapter and re-entry modules "to prevent unduedepletion of the spacecraft power supply. This is accomplished by means of anelectrical disconnect assembly in each module. These assemblies also containelectrical connections for monitoring spacecraft parameters during prelaunchcheckout. Each disconnect assembly consists of a plug and receptacle assemblyand a coupling device for securely retaining these components in the engagedposition. The disconnect assembly is capable of positive disengagement/separa-tion both electrically by a solenoid device and manually by a lanyard-initiatedmechanism. The re-entry module disconnect is normally released at approximatelyT-30 seconds. The adapter module disconnect is normally released at T+3 toT+5 seconds, with the backup device actuated by movement of the launch vehicle.A3.5.6 CONNECTORS.- Miniature bayonet (Bendix Pygmy) type connectors areused throughout the electrical subsystem except at external powerand interface connections. (See McDonnell Report 8612 for interface connectorspecifications.)A3.5.7 CIRCUIT BREAKERS.- The circuit breaker/switches used in thespacecraft are of the magnetic type. Choice of this type ispredicated on its relative insensitivity to thermal changes, which enhancesstability characteristics through temperature variations encountered in orbitalmissions. Hermetically-sealed integral switches are incorporated which providecapability of complete deactivation of selected circuits.A3.6A3.7

FABRICATION.- Same as basic specification.INSTALLATION.- Same as basic specification.

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AUAU.l

ENVIRONMENTAL CONDITIONSGENERAL.- This system and its components will be capable ofacceptable performance when subjected to the environmental con-ditions as specified in Tables I, II and III, as applicable.

ASA5.1A6A6.1A7A7.1A8A8.1A8.1.2loads.A8.1.2.1

QUALITY ASSURANCEGENERAL.- Same as basic specification.RELIABILITYGENERAL.- Same as basic specification.DATA REQUIREMENTSGENERAL.- Same as basic specification.TESTINGGENERAL.- Same as basic specification.ELECTRICAL POWER SYSTEM.- The electrical power system will betested with simulated dynamic loads and with spacecraft system

ENVIRONMENTAL TESTS.- Wire, fuses, circuit breakers, connectors,insulating materials, electrical panels and other electricalcomponents will be subjected in sufficient quantity to space environments,vibration shock, and acceleration representative of the Gemini mission to assureproper and adequate performance and operation.A9A9.1

DEFINITIONSGENERAL.- Same as basic specification.