project gemini familiarization manual spacecraft no 1

8/8/2019 Project Gemini Familiarization Manual Spacecraft No 1

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.- GEMINIROJECTEESCOPYDATE 18 _ J-96_R EV IB ED

PROJECT GEMINIFAMIL_IZATION MANUAL

FOR

sPAcsc_.:_:REPORT SEDIR :300 SERIAL NO. I

Mc,DONNELL __

SUBMITTED UNDER COI_"_R_.CT NO. NAS _-170

PREPARED BY _' ?t, APPROVED BY _/J///F_" _'/x-_._/

(_S I _ CREPRESENTATIVE


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INDEX OF EFFECTIVE PAGES

The pages of this Manual currently in effect arelisted below in nu-_rical order.

Page No. IssueA BasicB BasicC BasicD Basic

i-i Basic2-1 thru 2-16 Basic3-1 thru 3-4 Basic4-1 thru 4-3 Basic5-1 thru 5-2 Basic6-1 thru 6-35 Basic

A


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SEDR 3 00================================......................................................................................................................::.:.:.::.:._:..._::..._:,:,_._._,iiI ,!' i',iiNNOjECT,, ,, QEM

FOREWORD

The purpose of this _nual is to present, clearly and concisely, the descriptionand operation of the Gemini spacecraft systems and major components as installedin spacecraft number one. The pr_y usages of the _ual are as a familiari-zation-indoctrinetio_ aid, and as a ready reference for detailed _nfo_s_ion ona specific system or component.

The _-_uualis sectionalized by spacecraft systems or major assemblies. Generally_each section contains a system index illustration and a detailed word descripti_of the system, its operation and its _-_Jorcomponents.

The inforw_tion contained in this _nual is applicable only to spacecraft n-tuberone and is accurate as of iO July 1963. The m_nual will be revised periodicallyor as required by changes to the spacecraft.


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5E DR 300

II II I I III IIiiiiiii II I I I I I II I IIII

PAGESECTION ISPACECRAFTMISSION ......................................... l_l

SECTION IIMAJOR STRUCTURALASSE_IES ................................ 2-i

SECTION llICABIN INTERIORARRAN_ ................................. 3-1

SECTION IVENVIRO_2AL CO_W_ROLSYSTEM ............................... _-i

SECTIONVCOOLING SYSTEM ............................................. 5-1

SECTIONVlINS_ATIC_ AND CC_B_JNICATIONYSTEM ................... 6-1

C


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I_fRODUCTI_

Initiated by the NASA and _p.lemented by McDonnell Aircraft Corporation, ProjectGemini is the next logical step in the field of ms,ned space exploration. Closelyallied to Project Mercury in concept, and utilizing the knowledge gained fro_ theMercury flights, Project Ge_Ini will eventually orbit a two-._u spacecraft consi-derably more sophisticated than any employed so far.

The normally configured Gemini spacecraft is w_euverable within its orbit and iscapable of rendezvousing with and connecting to a second orbiting vehicle. Depend-ing upon the specific mission objective, it can stay in orbit up to fourteen days.Finally, upon re-entry, the re-entry portion of the spacecraft can be controlled ina relatively conventional landing.

Spacecraft number one, however, with which this _nual is specifically concerned,will perform a considerably less complex mission. Equipment requirements andperfon_nce capabilities are therefore relatively l_ted.

D


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SPACECRAFT MISSION

_iiiiiiiii_iiiiiiiiii_ TA B L E O F CO N T EN TS

...................... PAG.................... TITLE

.......................... DI_SCRIPTION 1 l.................... MISSIONo*o*_**.Ho_* .... ..... .......o_*_*_*_*_*_._**_._._........................... $PACI_CRAFT DESCRIPTION 1 1**o,o _ ..... . ........ ...... oo.o_HIHHHHH_IIIHHHI_.... o_o**o..... o_o,o,H**..... .o_,.. .......... . .......... o ..... .... ......... . . .. .. .. ........... ........ HIHHHHIHIHHHHHH..... . .. .. .. .. .. oHHHH_HHHHHHHHHHHHH_HH_IHH_HHHI_HHHHHHHHHHHHH. ................ HHHIH_HIIH_HH_HIHHHHHHHHHHHHHH_HHHHHIIHIHHH_HH;. .. .. . ... .. . .. .. . .. .. .. ........... ................... .. .. .. .. .. .. . . .. .. .. .. .. .. ............... .......

.. .. ... .. .. ... . , . .. .. .. .. .... ... .. ... ... .. ... ... ... ... .. ... .. .. ... ... .. ... .. ..

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

iiiiiiiiiiiiiiiiiiiiiigiiii_HHHH:H_HHH:_H


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_. " "_, i.:_i_" "

SECTI0_ I SPACECRAFT MISSION

__ION DESCRI_TIO]_

Gemini spacecraft n-m_er one will perform an unmanned orbital mission. The space-craft will not be separated from the launch vehicle. Recovery will not be attempted

The pr4_y objective of the mission will be to de_nstrate the structural integrity

of the spacecraft and launch vehicle through launch and a mln_,m of one orbit./Telemetered data eoncernin8 structural integrity, temperature, pressure and vibra-tin- will be obtained throughout the mission.

/SP ACE_ D ESC RIPT ION

Spacecraft number c_e conslb_s of production structure for the re-entry module andadapter. Within the spacecraft, the following equipment is accurately si_-_atedwith regard to size, shape, weight, center of gravity and method of attachment:

Co_mterIGS PlatformRCS Thrusters

/

OAMS Thrusters (except _ for_ firing thrusters)//

_11 other internal equlpcent is si_ted only with regard to _e_ight,c_nter ofgravity and method of attac_ut.


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MAJOR STRUCTURALASSEMBLIES

TABLE OF CONTENTS

- TITLE PAG::::::::.':::=:....._-:_""!_'.:=...:::.j_..._._iHii_iii_i_i_"_..:-_ GENERAL INFORMATION ................ 2-1!!iii!!}iiiii_i_}_ii'-.-':'..'=." RE-ENTRY MODULE .................... 2-1iliHiiiii!iiiiHiiiiiii!ii RENDEZVOUSAND RECOVERYSECTION...... 2-1iiiiiiiiiiii!iHi!iiiiHiii RE-ENTRYCONTROLSYSTEMSECTION ....... 2-7ii_iii_i!i!i!iiiiiiiiiiiii CABIN................................. 2-7iii_i!_!!!iiiiiiiiiiiiiii ADAPTER ............................... 2-12iiiiHi_iiiiiiiiiiiiiiiiiii RETROGRADESECTION..................... 2-14_i_iiii!ii!iiiiiiiiiiiiii ADAPTEREQUIPMENTSECTION............. 2-14i i i i i i i i ! i i i i i i i i i i i i i i i i i ii i i i i i i i i i i i i i ! i i i i i ! i i i i i i................ .... .......... ... ............... ................... ....................................

::ii::iiii:::_iii!i::::iliiiiiiii: : : : : : : : : : : :: : : : : : : : : : : : :: : : : : : : : : : : :: : : :


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" ; %!: !!

SECTION II MAJOR STRUCTURAL ASSEMBLIES

GENERAL INF ORM ATIONThe Gemini Spacecraft is basically of a conical configuration, (Figure 2-1), con-sisting of a re-entry module and an adapter as the two major assemblies. Space-craft construction is semi-monocoque utilizing titanium for the primary structure.It is designed to shield the cabin pressure vessel from excessive heating, noiseand meteorite penetration, (Figure 2-2). During spacecraft flight, the spacecraftadapter is aft with respect to flight path. See Figures 2-3 and 2-4 for space-craft orientation. TY is considered to be the top fo the spacecraft.

BE-ENTRY MODULE

The re-entry module (Figure 2-5) is separated into three primary sections whichinclude the rendezvous and recovery section (R & R), re-entry control systemsection (RCS) and the cabin section. Also incorporated in the re-entry moduleis the heat shield which is attached to the cabin, and a nose fairing which isthe forward section of the rendezvous and recovery section.

RENDEZV0_S AND RECOVERY SECTION

The rendezvous and recovery section (R & R), (Figure 2-5), the forward sectionof the spacecraft, is a truncated cone in shape and is attached to the re-entrycontrol system section With twmnty-four bolts. The R & R section utilizes rings,stringers, b,11_headsof titanium for its priw_ry structure and the external sur-face is composed of beryllium shingles.

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SEDR 300--. _ _ ,,"_! _:_ ._._!_i_:_:_:_._i_._:i!i_._ii i_!_!_:_i_!_iii_!_!_!:_:.':_!!!___ ._._i_.__:_.._i_:i_:_:__._ :_._.

'_ SPACECRAFT

ADAPTER _ _ RE-ENTRYMODULE

ADAPTERMATING ADAPTER RENDEZVOUSSECTON CABN

SECTION SECTION SECTION

RADAR

i

2RNGM AT ING LINERENDEZVOUS ANDRECOVERY SECTIONM AT ING LINE

R E - E NT R Y C ON TRO LSYSTEM SECTION/CABIN M AT ING LINE

RE-ENI"RY MODULE/ADAPTERMATING LINE

-- SPACECRAFT/LAUNCHVEHICLE MATING LINE

Figure 2-1 Spacecraft General Nomenclature FMGI-25

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S EDR 3 00*_..._.!.:::::::_:_:_::_:;.._:.:_:_:::_:_:_:::_:_::.:_::::_:_::::_:_::.::_.:::_:_::::_:::_:_i_i_i_:.__"_i:_::;:_.:;::::_:_:'_:_ " "':" "":"'"""::::""::"" :::_:..... ":'"":""""':"::::""":':""":":... .::.._!_!!:!:!_:..::_i .. ____LW_::_: "...!" x_

/ //

//

//

/ " /! \\ _ _-'\_ J\ j,,., //

Figure 2-2 Cabin Pressure Vessel FMG,-

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22667(LAUNCH AND Of'BIT CONFIGURATION)

120,00 DIA.

BYFigu re 2 -3 S pacecraft D im ensions FMGI-26

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_TOPITV

--._a t i,) J i

BY

Figure 2'-4 Stations Diagram FMGI-272-5


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iNGRESS-EGRESS

RE-ENTRY CONTROLSYSTEM SECTION

DOCKING BAR BASEHOLE AND PLUG

/--CABIN/ADAPTER __

RETAINING STRAP RENDEZVOUS ANDFAIRINGS _ RECOVERY SECTIONTYPICAL 3 PLACES)

OBSERVATION WINDOWSNOSE FAIRING

EQUIPMENT BAYACCESS

SMALL PRESSURE

DUMMY R.C.S. THRUSTCH

HEAT SHIELD

-E.C.S. EQUIPMENT DOOR

(TYPICAL 2 PLACES)DOCKING POINTS

(TYPICAL 3 PLACES)Figure 2-5 Re-entry Module StructureFMGI-2

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RE-EI_IRYCONTROL SYS_ SECTION

The re-entry control system (RCS) section is located between and mated to therendezvous and recovery section and cabin section of the spacecraft, (Figure2-5). This section is cylindrical in shape and is constructed of a center meg-nesium cylinder, eight strln_ers, two bulkheads and eight beryllium shingles forits outer skin.

CABIN

The cabin (Figure 2-_), similar in shape to a truncated cone, is mated to there-entry control system section and the adapter. The cabin has an internal

-_ pressure vessel shaped to provide an adequate crew station with a proper waterflotation attitude. This shape also allows space between it and the outer cc_i-cal shell for the installation of equipment. Structural design for the pressurevessel is 12.0 psi ultimate (burst) and S.O psi ultimate (collapsing).

Outlining the basic cabin and equip_nt bays is a heli-arc welded fr-_ sectionof ring segments, stringers and iongerons Employing titanium as a ma_or struc-tural _aterial, _h_ cabin incorporates unbeaded skins, seam welded to beadedouter skins and reinforced by longitudinal and vertical stringers on the surface.The outer circumferential surface is covered by Rene' _I shingles. Pressure b,,_k-heads are attached to each end of the cabin. On the top side of the cabin, tworegular hatches are provided of sufficient size to permit normal ingress and egressfor spacecraft checkout and instrumentation setup.

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j___._ SEDR 300

,J

0 0o o

_. z __- _ zo _z

Figure2-6 Access Doors (Sheet 1of 2) FMG

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//

Figure 2-6 Access Doors (Sheet 2 of 2) FMG,-

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SILL STRUCTURE _.

SI LL STRUCTURE

I LATCFI ASSEMBLY

(I'YPICAL 12 Pk_CES)--

LVIEW I_OTATED IS0 S_CTION A-A

Figure 2-7 Spacecraft Ingress/Egress Hatches EMG_ -

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E

Equipment Bays. The equi_ent bays are located outside of, and on the flat sidesof, the pressure vessel in the cabin, (Figure 2-5). Two bays are located c_ eachside of the pressure vessel and three bays on the bottc_ side of the cabin_ he-neath the pressure vessel floor. The bays are structurally designed for mountingof the equipment not requiring pressurizatic_ or ccm_onents having 8elf_ontainedpressurization.

Door_____s.To enclose the four equil_nent bays, two hinged, structural doors are pro-vided on each side of the cabin (Figure 2-6). These doors provide access to thecomponents installed in the equll_nentbays. The main land_-_ @ear _ays are en-closed by one door on each side of the cabin. The doors, inoperative on spaQe-craft number one, are attached to the structure by a total of i_9 bolts.

On the bottom of the cabin, equidistant from the centerline end _etween the lend-

ing gear doors, two additional doors are installed. The forward door allows ac-cess into the unpressurized compartment and the aft door provides access to theECS cc_tment which is a protion of the pressure vessel.

Hatches. Two large structural hatches, (Figure 2-7), are incorporated for sealingthe cabin ingress or egress openings. The hatches are s_trically spaced andare located on the top side of the cabin section. Each hatch is man,,nlly operatedby means of a handle and mechanical latching meachnism and is hinse_ on the out-board side. An external fitting is incorporate so a handle can be inserted forlatching and ,,nlarching by ground personnel.

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i IJ J I |OUT F.RGUkSS

PANE_

GASKET _,/.---GAS KETFP,Aht_ _ SRAME

--O'.RING

DETAIL A-A

H AT CH O UT ERML _EF)

OUTER WINDOW ASSY INNER ML _EF)OUT ER W IN DOW ASS EMBL Y

INNER WINDOW ASSYGLASS PANES

ZI_6.50

OBSERVATION WINDOW ASSEMBLY

MIDDL

INNER WINDOWASSEMBIYSTAT-O-SEAL

BOLT

_gure 2-8. Observation Window FMGt-_0

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I

The adapter, (Figure 2-1), is a truncated cone shaped, semi-monocoque structureconsisting of circumferential aluminum rings spliced together with extruded mag-nesium alloy stringers to form the basic structure. The forward end of the adapteris coupled to the aft end of the re-entry module by utilizing three titanium re-taining straps_ (Figure 2-9). The skin is formed of magnesium sheets which ereintegral with the magnesium coolant loops. These loops serve no function otherthan structural on spacecraft n,,m_erone.

RE_OGRAI_ SECTION

The retrograde section is the smaller diameter end of the adapter and is matedto the cabin section. An alm,_uum alloy "I" beam is used for structural rigid-ity and for mounting the d,,,..yretrograde rockets in this section.

ADAP_R E_IPM_ SECTION

The adapter equipment section is the larger diameter end of the adapter. Thissection utilizes al,,m_numalloy tubing in a truss beam arrangement for mountingof equipment

Spacecraft - Launch Vehicle Mating. The spacecraft is mated to the Titan IIlaunch vehicle through a continuous, machined al-mlnum alloy ring, (Figure 2-10).This ring, 120 inches in diameter3 mates with the launch vehicle mating ring.Twenty bolts secure the rings together. To provide for ali_-_nt, the launchvehicle incorporates one steel 3/16 inch diameter alignment pin and four indexmarks.

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.... SEDR300__:_a'_i_,}t__._._._i:__..J._.:::i._.,_:';.::i::_i._!.L._,._.::i.#:::::...._.i..::_i::.i._._,.::_._ :'_i._.::_'_:_._._..':_:o.__,'_

%,

SHAPED CHARGE"'"-.,.- " ASSEMBLY (REF)/

/" INERT

(TITANIUM)

PICKUP\\.SPHERICAL WASHER "_..SPACER "\_%

ii iiiii ...._ --%FAIRING

NUT ............. i.'" ..... S (REP)(REP) \_ "-.... i ,.---.Q............. ! - . .....A / "--_ " "

...._" STRAP(TITANIUM) -_"_-" ""'_... "" ":-, "-._;

'_'" \ RE-ENTRY MODULE"

//_(FAIRING (REF)REP)

J i i" STRAPSSEMBLY

SHAPED CHARGE

R EE N , R Y M O D O L E J. . . . . . . . . . . . . . . . . . . . . . . . .STRAPASSEMBLY (REE)(TYP2 PLACES)

RETAINING STRAP ASSY

ADAPTER (REP)J__HEAT SHIELD (REF)J / IS_"CTIO N A-A

Figure 2-9 Re-entry Module-Adapter Retaining StrapsFM_I-19

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:_ SEDR300

:__* " "i_._ ii _ ._!i'_.,..ii__l___ }i _ "I II I BB it ii

// OXI01ZER TANK

(REF)

QUAD 2 _\QUAD 4 / _'

QUAD 1 _. VEHICLE/

--E. "Z 2-- " _/- ;- y__ SPACECRAFT TO

LAUNCH VEHICLE

ATTACHMENT BOLTHOLES

T

FIBERGLASS FAIRIN G-_TIONMATING SECTION

SEPARATION ASSEMBLY !'bX LXJ

sj QDAD 1 AO 2SPACECRAFT

- ..... M/_TING LiNE oLAUNCHVEHICLE

RING ATTACH --BOLT BY

SECTION A-A(TYPICAL 20 PLACES)

Figure 2-10 Spacecraft Laurrch Vehicle Mating FMa_-,S


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CABIN INTERIORARRANGEMENT

TABLE OF CONTENTS

TITLE PA:::::""==::"="=-- GENERAL 3 1iiiiiiii_i}i:-'Hiii_._'_"_ PALLETS 3-1_!_ff_ff_ff_F_ ................................!iiiiiiiiiiiiiiiiii:_ili!._ STATIC SYSTEM .......................... 3-1i_iiiiiiiiiiiiiii_iiiiiiiii PITOT SYSTEM 3-4:::::::::::::: ::::::::::::: .......................i ! i i ! i i } i i i i i } i i i i i i i i } i } !i i i i i i i i ! i i i i ! i ! i i i i i i i i i H_ i i i _ i i i i ! _ i i i i i i ! i i i i i i i:::::::::::::::::::::::::::......................... o..... .. o .. .... o.........,.....,......o....,,.o......,...o, ................. ....... .......... .. .. .. .. .. . .. .. .. .. .. .. .................. . .............. ...... ,. ..... o ...........o.....o.,,,.. ........................... .,o.oo:::::::::::::::: :::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::

::::::::::::::::::::::::::::::::::::::::::::::::::::::


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G SEDR300

SECTION III CABIN I_TERIOR ARRANGE_

The instrumentation paJ_ets, pressure transducers, temperature sensors andacceler_neters are installed in the cabin. The instrumentation equipmentwill allow flight data to _e telemetered to the ground station. The cabinwill be purged and sealed prior to launch.

p_T.TRTS

The left and right pallets (Figure 3-1) are essentially platforms on which in-strumentation and oowm_nioation equipment is mounted. The pallets and a bal-last assembly are installed on the seat rails, which are located on the largepressure bulkhead. The pallets are restrained fr_ an upward or downward motionby a single point attac_-_nt located on the floor of the cabin beneath the pal-lets

S TA TI C S YS _

The static system is used for an altitude sensing source up to approximatelyi00,000 feet for the static transducers, located on the s_l pressure bulk-

head in the cabin. The atmospheric pressure is sensed by four static ports(Figure 3-2), sy_netricall spaced around the forward end of the conical sectionof the spacecraft. Through telemetry, the altitude is transmitted to a groundstation.

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RIGHT I -LEFT iNSTRUMENTATION PALLET

OUTER MOLD LINECABIN PURGEFITTINGS

Figure 3-1. Cabin Equipment EMGN

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DETAIL A.-A

SMALL PRESSURE __1

BULKHEAD \\DETAIL B-B\ -

\ -

/ i//STATIC PORTSTYPICAL 4 pLACES)-

Figure 3-2 Static and Pitot System FMGI-28

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The pitot system (Figure 3-2) is not used for this particular mfssicm. Thesystem is installed but will be plugged at the small pressure bulkhead in thecabin.

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ENVIRONMENTALCONTROL SYSTEM

TABLE OF CONTENTS

TTLE PAGii[ii::_._._::_2_'_:__ SYSTEM DESCRIPTION .................. 4-1!ii_!ii_...:i:#-_!i_i_ SYSTEM OPERATION ................... 4-1iHiiiiiiiiii_!_i_'_-_ SYSTEM UNITS ........................ 4-2

. .. .. .. ,_,_+..

......... ...... ..,,_. .. .. . . .. .. . . . .,. ,.

.... .....................i i ! i i i i _ i i i i i i i i i i i i i i i i i i li i i i i i g i i i i i g _ i i i i i i i i i i i ! i.......................... .................... ............................ .... ... .. ... ... ... .... .. ... ..

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


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I

SECTI_ IV ENVIROR_ENTAL CONTROL S_

SI_ rR.SCRIPTION

The Envlrom-_+ntal Comtrol System (ECS), as reflected in spacecraft c_e func-tions cml to establish and maintain a differential pressure of approximately 5.5psl between the interior of the cabin and ambient atmosphere. _s pressure isrequired during launch to prevent structural da-mge to the spacecraft. No pro-vlslcn is made to replenish or to clvoulate the gaseous contents of the cabin.

The differential pressure is established as the spacecraft ascends, reaching thespecified pressure at appraximately 25,000 feet. The pressure is established bythe aetiun of the cabin pressure relief valve.

During prepsratious for launch, the cabin is sealed with cabin pressure equal to-m_lent atmospheric pressure. The cabin is then purged with oxygen. During ini-tial spacecraft ascent, cabin pressure is automatically relieved to match the de-creasing atmospheric pressure. _hen the spacecraft reaches approximately 25,000feet_ however, where atmospheric pressure is apprc_w_tely 5-5 psia, the cabinpressure relief valve closes. Except for normal leakage, that pressure is main-tained throughout the reminder of the _ssiou.

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CABIN PRESSURE REr.'r_' VALVE

(seeFigure4-1)_e cabin pressure relief valve c_tains duplicate circuits for all i_nctic_s.Each of the duplicate circuits has a spring-loaded-closed poppet controlled bya servo element. This servo element controls a spring-loaded metering valvewhich determines the pressure within a diaphra_ chaz_r behind the poppet, thuscontrolling poppet positlo_. A s,_11 inlet bleed orifice aR_ts cabin pressureto the diephra_ c_-_er at all t_s. After the poppet opens, a larger orificeto the cabin is uncovered thus ensuring quick closing of the poppet.

CABIN _ FIT_NGS

An inlet and an outlet fitt_ is installed in the lo_er portion of the ECS com-_nt. These fittlnKs are used to purge the spacecraft cabin prior to launch.The fittings are self-sealing.

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I___ , SEDR 300':_'._4_.._."._._:. : ,:"_!: ;i ':._i_._i_.::_!. :_ :.

ICABIN PRESSURERELIEF VALVE

SPRING (TYP}SENSING

M ET ER IN G V AL VE(TYP) _ NSING CHAMBER (TYP)

CABIN PRESSURE(TYP)AND FILTER (TyP)

CHAMBER (TYP)

[ABIN OIAPHRAGM (EYP)

SERVO ELEMENT (TYP}' \-- POPPET VALVE

POP P ET VA LVE

I ABIN

I AMBIENT

SCREEN ASSEMBLY MANUAL SHUTOFF VAL VE

Figure 4-1. Cabin Pressure Relief Valve

FMG]-3


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COOLING SYSTEM

TABLE OF CONTENTS_oo+oo+++_::::::

ii__-_7-__ -_::_ _Z. TITLE PAGiiiiiiiiiiii._iii:._._-_-i SYSTEM DESCRIPTION .................. 5-1i _ i i i i i i i i i i i i i i i i i i i i i i qi i i i i i i i i i _ i _ i _ i i i i i i i i i i i i

iiiiii!ii_!!_iiiiii_H_iii


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_ __!_!!_!_ii i_i_i_i_!_i_!_!_liii :_:_ _ _i_i__i_i__i_i_Ii___!_

SECTION COOLINGYSTEM

SYS_ E_SCRIPTION

(See Figure 5-i)The cooling system consists of a series of coldplates installed on the instru-mentation pallets in the cabin. Equipment reT,_ing cooling is mounted direct-ly on the coldplates. During launch preparation, coolant is circulated throughthe coldplates from an external source to obtain maximum pre-cooling. At launch,the external source of coolant is disconnected. Cooling is then obtained only bytransfer of heat from the equipment to the pre-cooled coldplates.

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Figure 5-1. Spacecraft Cooling System EMG_-

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INSTRUMENTATION ANDCOMMUNICATIONSYSTEM

TABLE OF CONTENTS

"--"=::_........"-" TITLE PAGo.o**_ _**_**_iiiiiiiii_!i__.-.'-'#a_-i SYSTEM DESCRIPTION .................. 6-1!iiiiiiiiiiiiiiiiiiiiiiii-i SYSTEM OPERATION ................... 6-7ooooooo..oo.ooo:::::::::::::::::::::::::::iiiiiiii_ii_iiiii_iiii_ii_ SYSTEM UNITS ........................ 6-8! ! i i i i i i i i i i i i i i i i i i i i i i i i io....ooo.ooi i H i i i i i i i ! i i i i i i i i ! i i i i i ii i i _ i i i i i i i i i i _ i _ i i ! i i i i ! _i i i i i i i i i i i ! i i H i i ! i i i i i i i li i i i i i i i i i i i i i i i i i i i i i i i ! i i.o....... ..................... .. ................. ............................... ..................... .............................................. .._..... ... .. .. . .. .. .. .. .. .


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SECTICR VI INS_TION AND CC_ICATION

S_ _SCRIPTION

The Instrumentation and C_ication System (Flgures 6-1 and 6-2) serves asthe only co._=u_cation link between the ground and the Gemini Spacecraft. Ablock dia6ram of the Instrumentation and C..m-_nicationSystem is shown inFigure 6-3, and the location of the components in the spacecraft is shownin Figures 6-1 and 6-2.

The system-my be divided into three sub-systems: Communication, Instrumenta-tion, and Electrical Power. The Cv.._._ieationsub-system, consisting of a C-Band Radar Beacon, Phase Shifter, DC-AC Inverter, three C-Bend Ante---e, threeTelemetry Transmitters, _Au_Iriplexer,-nd a UHF Stub Antenna, serves as an in-stx--._ntationsignal ,,d spacecraft position data l_n_ between the spacecraftand ground.

The C-Band Radar Beacon, with its Power Divider, Phase Shifter, Inverter, andAntennas, facilitates ground radar tracking of the spacecraft during its entiremission.

The Telemetry Transw_tters, in conjunction with the _A_driple_er and the U_Stub Antenna provide a means of transmitting the data monitored by the instru-mentation sub-system to the ground station for evaluation and analysis.

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" _i_i::!_ _. ."':::: :! :::::"_:_!. ._"--:_ _::_ :,_

Figure 6-1. Instrumentation and Comm_ _ication System FMGI-1

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TELEMETRY TRANSMITTER(MID FREQUENCY)-

DC-AC(HIGH FREQUENCY)

..':;:_7;:";S;'."c"_AND._.DAR ,, ij'"_jJ ;_Mtt t! (LOWFRQUENCY

/'f

I Z /;..'f ,

e i

/DC-DC CONVERTER NOo

//RELAy "'PANEL /

SWITCH AND CIRCUIT BATtlERYBREAKER PANEL

Figure 6-2 Spacecraft Left Pallet (Sheet 1 of 2) eM

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SPECTRUMANALYZER(SYSTEM NO1

HIGH LEVEl. SPECTRUM ANALYZERC (SYSTEMO.

L OW L EV El .

.\) .

\

INSTRUMENTATION _1ASSEMBLY NO 1

INSTRUMENTATIONASSEMBLY NO 2

ASSEMBLYNO3" \x\\ I_V_OCOUPLE RBEE_:NC

JUNCTON\PACKAGE NO I

Figure 6-2 Spacecraft Right Pallet (Sheet 2 of 2) FMG

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S EDR 3 00____*_:_:._,._::::_ _ :::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::_ _:_'_._--'_..,._._:_::_, _,_"_

z_

- _,_-0

r , 1TI__a _ ,--a -Z o _ _ _oZ

Z_ _ Z_o z_? z>__ z_o__0 _0.... _ _


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_he Znstz,-_%ion su'_-syste:, con_)_.n_ _w._at;u:e sensors, In'essuze _-ens-(lucers, _11)_ti(m pick-ups, a_e].e _z_m_ters, an(1 microphones, In'ovt(les a meansOf _o_Litorlngthe physical co_itio_s and reactions of the spacecraft. _esesignals are fed to various Voltage Controlled Oscillators (VCO) which =o_ulatethe tele_try transmitters that transmit the obtai_ data to a suitablestatic. _he instru_enta%ion pick-up _evices are located throushout the space-craft.

_e Electrical Power sub-system _asic_ly consists of _e _i_ _atter_, severalcontrol relays an_ interconnecting wiring. The electrical power sub-syste_ sup-plies _r _irectly, o_ indirectly, for _ instrumentation an_ c_,-Icationc_nts.

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SEDR300 ._._*_._}_._i_i!!_i_i_i_i_ii_i_i_:_::_:_:_i!_ii_i_:..:_!}_`_.._' ..,.. ._

S_S_ OPERA_O_

The Instrumentation and C_cation System is _uall_ actuated prior to launch.

From the beginning of the Pre-Lau_ch Phase of the spacecraft mission through um-bilical release the lustx,r_ntation and Communication System may be controlledfr_a the blockhouse. Capability exists: to turn the telemetry transmittersand the C-Band Radar _eacon on and off; to apply main power either fro_ thespacecraft's win battery or from the blockhouse, and to monitor C-_a_d _eaconi_put po_r, bus power, and battery current. The blockhouse control and _o_i-tar feature is _de available by the Re-Entry module umbilical connector andvarious power c_trol relays in the spacecraft.

Prior to sealing the h-tch al_ circuit breakers are placed in the on position.Prior to lift-off power is applied from the blockhouse and the C-_ Radar_eacon and telemetry transmitters are turned on. At -m_ilical release thepower source is switched from blockhouse power to spacecraft _in battery power._ata c_er_i_ spacecraft temperatures, pressures, vibrations and accelerations,c_ti_uously _itore_ by the i_stz,_-_ntati_ system, is transmitted throughoutthe spacecraft's mission.

entire system, i_stx,,_ntati_ and cr_._...t_icatlons,ill be operat_-_ through-out the entire misslo_ of the spacecraft or as l_ted by the _attery life. Pri-

importance of data acquisition and trac_-_ capability is placed on the per-i_ starti_ with lauach and ending with i_serti_ of the spacecraft into its

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SEDR 300%.

. SYS_ U NITS

CO_RUNICATION

C-Band Radiatin_ Elements. Purpose: The C-_ Radiating R3ements are used toprovide radiation coverage and reception for the C-Band Radar Beacon. The an-tennas are located at the swatI end of the conical section of the spacecraftforward of the small pressure bulkhead.

Physical Characteristics: The physical representation of the C-Band RadiatingElements and their approx_te location in the spacecraft is sho_n in Fi@ure 6-4.The radiating ele=ents are mounted flush with the outside skin of the spacecraftand spaced 120 apart. The antenna unit is appr_imately 2.18 inches lc_g andhas a maximm vertical dimension of 3.41 inches. Each element has a TNC connec-tor for connecting the antennas to the phase shifter and power divider. Eachradiating element weiEb_ approx_w_tely 3.5 ounces.

Electrical Characteristics: The C-Band Radiating Elements receive a properlycoded signal from a ground tracking station which, when found acceptable by theC-Band Radar Beacon_ triggers the transmitter to reply with a _ kilowatt peakpulse modulated signal at a frequency of 5765 MC by means of the C-Bana Radiat-ing Elements. The combination of the three antenna elements gives a radiationpattern which extends in a1_ directions except forward and aft of the SpaceCTCft.

Phase Shifter. Purpose: The Phase Shifter compensates for areas of low_ or noradiation coverage het_en lobes of the three C-__nd Ant_---_ radiation patterns.The Phase Shifter, Power Divider, and the three radiating elements cc_prise an


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_'_-_._'_._._i!_.i.".._#_::_ _ :_!_i!ii::_::_:_::_:_:::::::_:::_::!_i_ii_!..`.!_!_._.._.:.%_"_":_'_:::::_i::"": :" _ _i_:._._,,_,_,,___.:..i::::-..:................':'._..'.......'i.......i"i.,_,_i_i!,__ii.._..._-:.

Figure 6-4 C Band Radiating Elements _M-

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SEDR300

antenna system which provides essentially complete coverage for all orbitalattitudes.

Physical Characteristics: The physical representation and approximate locationof the Phase Shifter is shown in Figure 6-5. The phase shifter is located onthe small Pressure b,1_head outside the cabin and has a total length of approxi-mately 4 3/4 inches and a diameter at the small end of about 1 1/2 inches. Theph_e shifter contains three connectors to connect the phase shifter to the powerdivider, to a C-Band Radiating Element, and to a DC to AC inverter. The phaseshifter weighs apprc_tely 20 ounces.

Electrical Characteristics: The sehe-_tic of the pbA-e shifter is shown inFigure 6-5. A 26 VAC 400 eps input to the phase shlfter, obtained from a DC-ACinverter, is half wave rectified and applied across a coil wound around aferrite material. Due to the characteristics of the ferrite material, the RFsignal from the power divider is delayed 0 to 180 +-200 at the rate of _00cycles per second. The changing phase shift of the RF power on one of the C-Band Radiating Elements with respect to the other two shifts the lobe of thatantenna by approxlwately + 45 degrees; thus, giving the effect of an -1-_stideal circular radiation pattern around the longitudinal axis of the spacecraft.

DC to AC Inverter. Purpose: The DC to AC Inverter is used to power the C-BandPhase Shifter.

Physical Characteristics: The DC to AC Inverter is shown in Figure 6-6. Theinverter is approx1_tely 1.7 incheshigh,1.5 inches wide, and 3 inches i_.

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_'O C BANDRADIAIINGELEMENt

OSTEP-UP _

2.5:1

T O PO WERDIVIDER

Figure 6-5 Phase Shifter FMG1-2

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SEDR 300

I

, ,0R4 i RS RYEL 1"1 BRN _- B ORN _ RI 12 RED _ BLU

IO PHASE I R3 G GRN 13 REDSHIFTERYEL_,, BRN _._ :_ c_T _ _ c_BRN -

R6 RT

i II Q4I

,_ _ YEL CR4 JI(

I BLU / R8 RII RI2

BENDIXPT1M-I 0-6P(101)

Figure 6-6 DC-AC Inverter FMG1-5

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SEDR 300

The inverter is _unted on the left hand pallet of the spacecraft. One s_ngleconnector is used for both input and output vol_es. The weight of the inver-ter is appro_mtely 0.6 pounds.

Electrical Characteristics: A DC input voltage is supplied to the inverter(Figure 6-6) frc_ the spacecraft main battery. The inverter provides a 26VAC, 400 cps output required for operation of _ phase shifter. The inverteris designed to work into a capacitive load; therefore, a two _icroforad capaci-tor is placed across the output terminals of the inverter to co_ensate for theinductive input of the phase shifter.

Power Divider. Purpose: The Power Divider is used to give equal power distri-bution to the three C-Baud Radiating Elements.

Physical Characteristics: The physical representation of the power divider isshown in Figure 6-7. The paver divider is located on the small pressure bulk-head outside the cabin and measures fro_ the left side antenna connector to thebeacon co_ncector approx_tely 3.8 inches. _he distance between the twoscrews is approx4_tely 4 inches. The power divider coatai_s two t,m4,_ ad_ust-_nts, and 4 connectors of which one connects to the C-__nd Beacon, two connectto the left and top antenna, and the fourth oue c_m_eets to the phase shifter

that is in series with the right a_te_na. The power divider weighs appr_ctmately6._ pounds.Electrical Characteristics: The power divider is basiea_ly a cavity type powersplitter. Dur_n_ beacon transmission the power is delivered to the power divider.

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$EDR300

TO PHASE SHIF1BR(RIGHT ANTENNA

TUNING STUB CAPS

RADAR BEACON

Figure 6-7 Power Divider PMG_-9

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S EDR 3 00......_4 ._.%._:_:_:_:_:_:._:':i:_::_:_::::::_:::_'_:!:'_::::':::::_8:::::.::::::::: :_::_ii:__ !_ !!: !_ _: :__i:_ '_ _.... ......................

ANTENNA

P EA K P OW ER

RECEIVER_ RECEIVER TEST POINTSDIRECTIONAL PEAK POWER -- TRANSMITTERCOUPLER AMPLIFIER PRF

@ MONI ToRPEAKOWER I

O TRANSMITTER PRE\

TRANSMITTER AND LOCKOUT OVER [NTERRO-OATION

@ TRANSMITTER OVER '_FI_QUENCY INTERROGATIONIIi ,

CIRCULATOR MIXER I.F. -_ AMP DECODER DELAY COINCIDENCE DELAY(DIPL_XER)

IIRCULATORPRE-SELECTOR (MIXER) IMMUNITY@ l

RECEIVERFREQUENCY

LOCAL VARIATION VARIATIONOSC. CORRECTION CORRECTION

_CE_VER _; ,%FREQUENCY SLOPE #] SLOPE #2LINE POWER VARIOUS_l FILTER SUPPLY VOLTAGES

Figure 6-9 "C" Band Radar Beacon Block Diagram FMGI-

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SEDR 300._::.._.:..'_ ._ ........... ::::::::.......:.......:........:::::..,..:..... .:.. ."__ : _.

coupler to one half of a dual ferrite circulator. The characteristics of the

ferrite circulator are such as to isolate the transmitter from the receiver;thus, providing the capability of utilizing a single antenna, or antenna systemfor both reception and transmission. The beacon utilizes a superheterodyne re-ceiver to provide the required reception. The receiver is tunable by maans of athree stage pre-selector over a range of 5600 MC to 5800 _C. The assigned re-ceiver center frequency is 5690 MC. The output of the pre-selector is combinedwith the local oscillator frequency in the crystal wd_er which pro_ucee an o_t-put IF frequency of 80 MC. The local oscillator is of the metal-ceramic triodecavity type and the _xer contains a ferrite circulator for isolation _et-_een thelocal oscillator, mixer, and pre-selector. The output of the mixer is then amp-lified by three tuned IF amplifier stages followed by a video detector an_ avideo pre--mD_lifier. Additional amplification is obtained by a pulse =-_.lifier.The purpose of the decoder is to initiate triggering of the transmitter after acorrectly coded signal has been recelved. The system delay in co, unction withthe delay variation correction circuitry provides for a constant fixed delay use_in determinin_ the exaot position of the spacecraft. The beacon incorporates aCW immunity circuit that prevents the transmitter fr_ being triggere_ by rand_noise. The noise level is reduced below the triggerin_ level of the transmitterby controlling the gain of the pulse amplifier. The tran-_itter utilizes a mag-netron and provides a One kilowatt peak l_,l,emodulated signal at a frequency of576_ MC to the power divider. The beacon is powered by a I)C-DCconverter employ-ing a magnetic amplifier and silicon controlle_ rectifiers. _e converter pro-vides voltage regulation for input voltage variations between 18 and 32.5 VDC.

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$EDR 300

_ _7__.:._._._. _'+_::::::::::: ::::: _:.:::::_::::..._:::::::::::::::::::::::::::::::::_:::_::::_:::!:!:::_:::::_:_:_:_.._:_._._._._ _ __ _ _ __k_ __,=_,_

The input to the converter is filtered by a pi-type filter to mi_Ize any linevoltage disturbances.

Tele_etr_ Transmitters. Purpose: The three telemetry transmitters provide aradio frequency (RF) link from the spacecraft to ground co-H-_nicationfacili-ties for transmission of various data obtained by the instrumentation sub-system.

Physical Characteristics : The three telemetry transmitters are identical ex-

cept for the operating frequency. The physical construction and approx_-_telocation of the transmitters in the spacecraft is shown in Figure 6-10. Thetransmitters are approximately 2.75 inches high, 2.25 inches wide, and 6.5inches long. The three trs_smitters are mounted on the left cabin pallet.Each transmitter contains three connectors : a DC power connector, a RF out-put power connector, and a video input power co_ector. Each transmitter weighsapproximately _I ounces.

Electrical Characterist_cs : The three FM transmitters are frequency modulatedand operate continuously to provide an output po_ _"of a minimum of two watts.The low frequency tran_tter, operating at 2_0._ MC, transmits data concerningvibration, acceleration, acoustic noise, pressure, and scanner fairing break-wire continuity. Temperature and vibration data is tran_Itted by the mid fre-

quency transmitter at 2_6.S _. The high frequency transmitter transmits dataconcerning vibration, te_perature_ acoustic noise, pressure and scanner fairin_brea_-wire continuity at a frequency of 259.7 _C.

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,,:._.:.:,:,,,:.,,:.:.:,:,:,: .............. . .... , . _:.:.: ,,:

Figure 6-10 Telemetry TransmitterFMGI-]4

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I Jl I

UPIFQuadri_lexer. The UHF Quadriplexer provides isolatic_ between three tele-metry transmitters operating simultaneously into a c_ antenna.

Physical Characteristics: The physical representation and approximate locationof the quadriplexer is shown in Figure 6-11. The quadriplex_r is mounted on theleft cabin pellet and is approximately 6.7 inches wide, _.12 inches high, and5 inches deep. The quadriplexer contains five connectors: the UHF stub antennaconnects to JS, the mid frequency, low frequency, and _gh frequency telemetrytransmitters connect to Jl, J2_ and J_ respectively. Connector J3 is not usedin the spacecraft number one configuration. The quadriplexer weighs approximatelythree pounds.

Electrical Characteristics: Figure 6-11 shows the schematic of the qua_riplexer.Basically, each channel consists of a high Q cavity tuned to the correspondingtransmitter frequency. All channels are isolated from each other without appre-ciably attenuating the RF signals passing throu_ it. Each channel has the capa-bility of being retuned in case the transmitter frequencies or the functio_ ofthe quadriplexer change. The outputs of the telemetry transmitters are fed intoJl, _2 and _ of the quadriplexer and brought out with a sin@le antenna connectorthus eliminating the need for an antenna for each transmitter.

UHF Stub Antenna. Purpose: The U_F Stub Antenna provides s_w_taneous radia-tion coverage for the three telemetry transmitters.

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SEDR 300__11111____i_-__ ;.. _ u

/tt_", _ L II O J_

MID TM

LO TM

J_ 0 [&RUHFSTUB ANTENNA

T I_ 0 J3/_ J _ , _

_ II oJ_

Figure 6-11 UHF Quadriplexer FMG,

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"_!i;_'_i.:.,:_ii_!__!::__:_ii_i_::___!!!!!!!!:i_i____ii:_i_i_i_:_i_ _!__'_:_:_ii! " "_(: : :!:i:_:_:_:i:!:!:i:_:!:_:!:_:_:_:_!_!_i_i:i:::!:3!;

Physical Characteristics: Figure 6-12 shows the UHF Stub Antenn- and its approxi-mate locatio_ o_ the nose of the spacecraft. The ante_, is mounted c_ the R & Rsection and extends about six inches forward from the spacecraft nose fairing.The antenna has an ann,_!ar slot in which a variable capacitor is _ounted and isapprox_tely 11.2 inches in length, 0.63 inches in diameter, and is for the majorportiou of the length hollow. The antenna cont-_s one connector to mate with theoutput of the quadriplexer.

Electrical Characteristics: The D_F Stub Antenna provides sim_taneous radia-tion coverage for the three telemetry transmitters. The v_riable capacitor placedacross the annular slot provides broad-band tuning capability of the antenna. Theradiation pattern of the antenna is identical to that of a q,,-rterwave stub.

INS_ATION

Te_rature Sensors. Purpose: Resistive element temperature sensors and ther-mocouples are provided to sense various spacecraft temperatures and to convertthese temperatures into proportional electrical signals.

Physical Characteristics: A basic thermocouple is shown in Figure 6-13. Althoughvarious physically different terminations are used, the basic construction consistsof two elements of which one is AI,-._I,and the other Chr_nel. The three differenttypes of resistive element temperature sensors are also shown in Figure 6-13. The0 - 300 F sensor measures_ excluding the mounting flange, approx_-_tely 0.4 incheshigh, 0.75 inches wide, and 1.5 inches long; the 0 - _00O F sensor is appr_x1._tely

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,4_, $EDR 300

PLUG

"IUNING ::

A ,_ ENNA MAS

PLUG

E PON R ESIN

qG FLANGE

SULATOR

Figure 6-12 UHF Stub Antenna IMGI-3

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I I I

0.5 inches hi_h, 0.5 inches _rlde, and 2 inches long; and the 0 - 600 F sensoris apprc_w_tely 0.3 inches high, 0.7_ inches wide, and 1.5 inches long. Figures6-13 _hru 6-15 identify and show the location of all thermocouples and resistiveelement temperature sensors. The sensing element of the resistive element tem-perature sensors are made of pure platinu_ wire encased in ceramic insulation.

Electrical Characteristics: _asic-_ly, the resistive element tez_erature pick-up consists of a bridge circuit which is energized by the +_ VDC regulated out-put of _._ DC-DC converters. One leg of the bridge is made of platinum wire.The characteristics of the platinum wire causes a proportional resistance changefor a c_ in temperature; thus, unbalancing the bridge and providing an out-put proportional to the change in temperature. Two resistive element temperaturese_so_s provide a reference temperature for all thermocouples by _itorlng thetemperature of the reference Junction.

Ther_oco_ples are used for measurements of temperatures exceeding I0000 F whererapid response to a large temperature change is necessary. The outputs of thether_ocouples and resistive element temperature sensors are 0 - 20 millivoltsand are supplied to the low level c_tator for sampling.

Pressure Transducers. Purpose: Two types of pressure transducers are providedto sense various spacecraft pressures and convert these pressures into propor-ti_l electrical outputs.

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Physical Characteristics: The pressure transducers ere shown in Figure 6-13.The transducers are appr_x4mtely 1.25 inches high, 1.25 inches wide, 1.75inches long. The mechanical portion of the transducer is a bellows which isphysically attached to a wiper of a potentic_eter. The transducer weighs ap-prcad_ately O.31 pounds. Figures 6-13 thru 6-15 with their corresponding chartsshow the ne_e and location of all pressure transducers.

Electrical Characteristics: The pressure transducer potenti_eters are poweredby the +_ VXX_regulated output of the DC-DC converters. Basically, a change inpressure causes the bellows either to contract or expand by a proportional a_ntwhich moves the _riperof the potentio_eter that is physically connected to thebellows; thus, providing an ontput voltage proportional to a pressure chan_.The derived output ranges fr_n 0 - 5 VDC and is supplied to the high level cc_-mltator input for s--_14-_. The ran@es of the pressure transducers are 0 - 15PSlAand 0 - 6 _S_.

Vibration .Pick-U_s. Purpose: Force _alance servo accelero_eters and crystaltype accelerc_ters are provided to monitor various structural spacecraft vi-bratic_s and to convert these vibrations into proportional electrical signals.

Physical Characteristics: Low and high frequency vibration pick-ups are provided.The low vibratio_ pick-up consists of a force balance servo acceler_neter end afilter shown in Figure 6-13. The accelerometer is appr_Imately 3 inches lon_,1.2 _nches wide, and 1.2 inches high; and the filter bias unit is approximately1.8 inches high, I.I inches wide, and 2 inches loz_. The high vibration pick-ups consist of a crystal type acceler_neter followed by an a_plifier. Both

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.... SEDR 300glgW__==============================================================================================:.::i :.:::: .:::: ::::::::i:::::i:!:i:_:!:!:!:i:!:!_i_:!:i_i_i!_i!!!i_!i_`?`_._i_`__`_-`_'_''_ _Ill I

units are shown in Figure 6-13. The accelercm_ter is 0.625 inches hexagon shapedat the bott_ and 0.785 inches high; and the a_pllfier is appro_-_tely 1 _nehhigh, i.I inches wide, and 2.5 inches ic_. The _eight of the acceler_meter isapprox1_tely 1.1 ounces, and the Amplifier weighs apprn_tely 3 ounces.

Electrical Characteristics: Low frequency vibrati_ measurements, ran_ from1 to 30 cps, are acco_lished by a force balance servo accelero--ter and a filter.

The acceler_ter together with the filter offers a frequency response from i to_0 cps with an _n_inite attenuatic_ at 0 cps. _ne accelerc_eter is esaential_ya torque balance closed loop system utilizing a pendulous mass supported by anextremely low frictice_Jmeel bearing. _asically, a elect_tic posltlc_ de-rector notes the slightest _ov_nt of the _ass, caused by a c_uge in velocityof the spacecraft, and applies a directly proportional electrical signal to aservo a_lifier. The Output of the servo a_li_ier is applied to a _r_e gen-erator which tends to restore the mass to its equilibrium position. B_ measur-i_ the amount of current required for the mass to reach its equilibrit_ position,the acceleration may be determined. Since the primary interest in this case islow frequency vlhratic_s, the output of the accelerometer is fed through a i to30 cps filter. The c_tput of each filter is applied to a VCO. The output ofthe accelercm_ter is 0 - 5 VDC, directly prol_ic_al to the a_llt_de of vibra-ti_. T_e mid (20 cps to 600 cps) and high frequency vibrations (up to 2000 cpe)are monitored by a crystal type accelerc_eter fo_ by an a_lifier.crystal accelerc_eter provides an electrical si_..1 Output proportional to the_mount of acceleration. _he output of the aecelerc_eter is then a_lified to a

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5EDR 300:_i!_!...............:"'""::!,::!:i"!'""_" ::' ii:_:::i"-:::'-": ::""_."':'"::'_':"i_'_-_:"

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sufficient amount to drive the various voltage controlled oscillators. A?] vi-

bratio_ pick-ups are pop,red by the DC-DC converters.

Accelerom_ters. Purpose: Three acceler_meters are provided to measure space-craft axis accelerations and to convert "g" forces into a proportional electri-cal signal.

Physical Characteristics: The physical characteristics of the acceler_m_tersare identical to the force balance servo accelerameters described under vibra-tion pick-ups.

Electrical Characteristics: The operating characteristics of these accelero-meters are identical to th_ force balance servo accelerometers described undervibration pick-ups except that the response is extended to 0 cps.

Acou st ic Noi se Pick - Ups. Purpose: Two microphones are ,itilized to pickup any

acoustic noise in the spacecraft cabin and to convert this noise into propor-tional electrical signals.

Physical ChAracteristics: The acoustic noise pickup system consists of twomlcrophones, 2 amplifiers, 2 low-pass filters, and 2 spectrum analyzers. Thespectrum analyzers are located on the right cabin p_llet of the spacecraft.For location of the other ccm_oonents refer to Figure 6-1.

Electrical Characteristics : Noise picked up by the microphones in the cabin isamplified and then fed to a spectrum analyzer. Each spectrum analyzer consists

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:::::::::::::::::::::::::::::...............:.- -:::::::::..........:................................::..........:..............:......._.._...._.:.._,,.I

of nine filters of which each passes an octave band of frequencies between 37.5cps and 9.6 Kcps. The outputs of the spectrum analyzer are converted to propor-tional 0 to 5 VDC signals and supplied to the high level co,.,,,_tatornput forsampling.

Scanner Fairin_ Break Wires. Purpose : Two break wires are utilized to _w__itorthe environmental cc_ditio_ of the horizon scanner scanner.

Physical Characteristics: This monitor, or sensing system, consists of two piecesof wire attached to the structure of the scanner fairing. The characteristics ofthe wires are such that they will break in the event of excessive elongation orvibration of the scanner fairing.

Electrical Characteristics : 0_e end of each wire is placed across an E_ of ap-prox1,_tely _.2 VDC obtained from a voltage divider In the Instrumentatic_ package.The other end of the wire is continuously sampled by the high level o.......tator.A broken wire will be detected by the high level co_._,-,tatorn the form of no in-put voltage.

DC-DC Converter. Purpose: The two DC-DC converters supply regulated DC voltagesto the instrumentation components and reference voltages to the high and low levelc_tators.

Physical Characteristics : Both DC-DC converters are identical physically and elec-trically. The conVerters are mounted o_ the left cabin pallet of the spacecraft.Each converter is approximately 5.5 inches high, 5.5 inches wide and T inches long.Each converter has a test cc_nector and a connector for input and output voltagesand weighs approx_w_tely 7 pounds.

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_,_ SEDR 300 _:._._._._i_!i!i!i!._i_i_i_:_i_!_i_:!i!i;i:_i!i!_i!i_::_::i:_:_:!i!:_i_i_!i_i_!_ii!i!_ii_!_ii_!_!_i_!_i_ii!_!_i_!_ii!_!_!_i._i_._:i_._i_ ......._i_l'_::_!_iifi!..: "i_iiii::_:::.i:::_:_:_ii::iii::i:i_ .:_:: _!i '::_i_''_ _II

Electrical Characteristics : The DC-DC converters are essentially voltage regu-lators. The converters operate on main bus voltage variations between 18 and_0._ VDC. Each converter provides three different regulated outputs, -2_ VDC_+_ VDC, and + 5 VDC. The DC-DC converters apply power to all instrumentationsub-system components.

C_,_tators. Purpose: A low and high level commutator provide a _eans of s-wDl-in_ various _-st_tati_n signals.

Physical Characteristics: Both c_._,tators are located on _? right cabinpallet of the spacecraft. The low level c_tator is appro_tely 3.9 incheshigh, 5 inches wide, and 5-9 inches long. The _gh level c_._!_._,ta_oreighs ap-pro_tely 2.9 pounds, and the low level c_,_._,_atoreighs abont 6.7 pounds.

Electrical Characteristics : Each co._.-,tatoris pulse amplitude modulated andhas the capability of _._Itoring 88 inputs at the rate of I i/_ samples persec_. _he low level commutator samples input signals between 0 and 20 _i_-volts; _ly, 63 th_ocouple outputs and 16 resistive element temperature sen-sor inputs. To _-_-_ze errors due to c_.._...._ode disturbances, the l_w levelc_,_,tator accepts a double ended input; _t is, the inputs are balanced to ac_ return and read the difference between the _h and low signal input wires.The high level c_,tator samples input signals between O _ 5 VDC, these are:pressure transducer signals, acoustic noise si_mls, scanner fairing break wiresi_mls, and main bus power. The high level commutator accepts a single endedinput_ that is, the input is referenced to a co_ return.

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SEDR300 ...y_/._._ =========================================================================:::::::::::::::::::::::::::::::::::::::::::

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Instrumentation Assemblies. Purpose: The instrumentation assemblies proviae anoutput frequency variatiun for a correslx_ding input voltage variation.

Physical Characteristics: In reference to Figure 6-1, the number 3 instrumen-tation assembly contains 7 volta6e controlled oscillator (VCO) modules; _e _r,and one hardline amplifier. Each VCOmodule is approximately 1.3 inches wide,1.6 inches high, and 0.8 inches deep. _he number 2 inst_---_ntation assembly con-tains _ VCO modules 3 o_e mixer I and one h-rdline amplifier. The number 1 instru-mentation assembly c_-e 9 VCO modules, one m4wer_ and o_e h-rdline amplifler.All three instz,,w_ntatlonassemblies are locatea on the right cabin pslletof the spacecraft. Each VCO m_ule weighs appr_4_-tely 2.2 ounces.

Electrical C_-racteristies: The _er 3 instrumentation assembly accepts 6 vi-_ratic_ pick-up signals3 and the output of the low level c_.,._-,,tator.ach of theinputs to the inst_-_ntation assembly is fed to a VC0. The VCO's provide an out-put frequency variation for a corresponding input voltage variation. The VCO,operat_-_ at an assigned center frequency deviates from this center frequency witha deviatio_ in input voltage. The outputs _f the VCO's are cc_bine_ in a m_er,the output of which is used to modulate the mid frequency telemetry transmitter.The outputs of the VCO's are also applied to a _-r_llne amplifier used during_round checkout of the spacecraft. The mm_er 2 instrumentation assembly receivesan ix_mt for each c_e cf its _ VCO modules frc_ the low level c_.-_tator, the highlevel c_tator, and 2 vibration pick-ups. The outputs of the VCO's after mlxin6modulates the high frequency telemetry transmitter. A har_line amplifier is also


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SEDR300

provided for the number 2 _nstr_ntatic_ assembly. The D,_er I ins_r_nentati_nasse_l_ receives its VCO _puts fro_ _ vibration pickoups, _ acceler_ters, andthe high level cow,tator. The m_ed VCO outputs _f the _,1_er 1 inst_-_-ntati_nassembly is used to modulate the low frequency telemetry transmitter; a hardlinea_plifier provides capability for ground checkout.

Electrical Power, Purpose: The Electrical Power Sub-System supplies power di-rectly, or indirectly to all instx-_-_ntationand co_mication system cn_!x_nents.

Physical Characteristics: The _aJor components of the electrical power sub-system,a battery, relay panel_ and switch and circuit breaker panel, are located on theleft cabin pallet of the spacecraft. The battery is approx_wtely 5.9 incheshigh, 6.38 inches wide, and 8.3 inches lo_g and weigh, appr_-_tely 17 pounds.

Electrical Characteristics : A _5 aw_ere_our 16 cell battery, having a n_-alterminal voltage of 24 VDC, serves as the only power source for the spacecraft.The switch and circuit breaker panel c_tains a main power switch used to arm,or disarm the power bus, and 7 circuit breakers to provide circuit protectio_for the DC-DC converter _i and _, Telemetry Transmitters _i, _ and _, the C-Baud Radar Beacon, and the DC-AC Inverter. The relay panel's main power re1_yand four control relays are utilized for ground control prior to launch. The_, power relay, when energized frc_ the blockhouse, disconnects the battery

project gemini familiarization manual spacecraft no 1

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