gemini design features

8/8/2019 Gemini Design Features

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GEMINI DESIGN FEATURESWilliam J . Blatz

Senior Project Engineer, GeminiMcDonnell Aircraft CorporationThe Gemini program, vhich vas init iated by th eNASA approximately 16 months ago, is being implemented by th e McDonnell Aircraft Corporation as th enext logical step in th e nation's manned spacecraftprogram. The underlying concept of th e Gemini de sign is to ut i l ize th e Project Mercury background toth e ful les t possible extent as a stepping stone to apractical operational spacecraft. The key vord hereis "operational." Project Mercury's basic objectiveva s simply to pu t man in space and bring him safelyback . Gemini, in contrast , aims a t exploring andexploiting man's abi l i ty to function in space and todevelop t ru ly operational systems and techniquesapplicable to a variety of missions. Retention of

the basic Mercury aerodynamic configuration and re entry heat protection concepts has permitted development to proceed vith a minimum of costly and time consuming f l ight demonstration testing . This hasalloved emphasis to be placed upon development ofth e various spacecraft systems . I t is in the l a t te rarea that th e rea l advances of Gemini over Mercuryare evident.Before going on to a more detailed descriptionof th e spacecraft and i t s systems, i t might be veIlto f i r s t examine th e Gemini mission objectives andconsider hov they have influenced the spacecraft de sign . Quoting from th e Gemini contract :"The objective of this contract is the research

and development of a versatile general purposespacecraft fo r th e accomplishment of space missionsof increasing complexity .

"Specific objectives are : (not in order of im portance)a. Fourteen -day earth orbi ta l f l ights .b. Controlled land landing as primary recoverymode.c. Demonstrate rendezvous and docking vith atarget vehicle in earth orbit as an oper at ional technique.d. Develop simplified spacecraft countdovntechniques and procedures fo r accomplish ing th e rendezvous mission vhich ar e com

patible vith spacecraft launch vehiCle andtarget vehicle performance .e. Determine man's performance capabilit iesin a space environment during extended

missions . "Addi t ional major design considerations vereth e designation of the Titan I I as a launch ve hicle, th e selection of th e Agena as a rendezvousand docking target vehicle, and th e requirement fo ra tv a-man crev.

Each of th e stated mission objectives repre sents a significant step forvard, and each dictatesspecific design requirements beyond those imposedby th e Mercury mission . In Table 1, an attempt ismade to categorize according to miSSion objectiveth e many Gemini subsystems and design featuresvhich are nev or significantly improved over co r responding features in th e Mercury spacecraft .Consider f i rs t th e 14 -day mission . This re quires an order of magnitude improvement in meantime -before -failure in many areas to achieve comparable mission rel iabil i ty to th e Mercury program.Added electrical energy requirements have dictated

th e selection of a fuel cel l system for pover vhilein orbit. Cryogenic storage of hydrogen and oxygenfuel cel l reactants and breathing oxygen are usedto conserve veight and volume . Heat generated byeqUipment and crev is rejected to space by a radi ator using a liquid coolant vhich is circulatedthrough cold plates and heat exchangers in place ofth e vater boiling technique used fo r th e shorterMercury missions. Pulse code modulated telemetrygives th e high data transmission rates needed todump th e stored information during the limited timeover ground tracking stations. Due to th e bulk ofexpendable supplies, the adapter betveen th e launchvehiCle and the spacecraft re -entry module is ut i l ized as an equipment compartment and is retainedvith the spacecraft in orbit . I t is jett isonedjus t prior to re-entry in contrast to th e Mercuryprocedure of leaving th e adapter attached to th elaunch vehicle.

Achievement of th e second objective of a cont ro l led land landing a t a pre-selected point in volves control of both th e re -entry trajectory andth e f ina l touchdovn maneuvers. Elf offsetting th ecenter of gravity of th e re -entry module approximately 1 . 75 inches from th e longitudinal centerl ine , an aerodynamic l i f t - to-dxag ratio of approximately . 22 is generated . The resulting l i f t vectoris directed as needed to modulate th e re -entrytrajectory by controlling th e ro l l att i tude of thespacecraft. The ro l l att i tude is adjusted inresponse to error signals generated by an on-boardinert ial guidance system consisting primarily of aninert ial measuring unit and a general purposedigi ta l computer. The iner t ia l guidance systemalso performs orbit navigation functions to keeptrack of present position and compute th e properretrograde time to allov touchdovn a t any preselected si te vithin th e maneuvering capability ofth e vehicle . A digi ta l command receiver permitsperiodiC up-dating from ground tracking stations.ContrOlled landing is accomplished by means of aparaglider vith final touchdovn on a 3 -skid landinggear. Ejection seats are provided as a backup fo rth e paraglider. They also serve as a crev escapesystem during th e early portion of th e launch andpr e - launch mission phases.


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The third mission objective is to rendezvousand tben to dock witb an Agena target vebicle.Target bearing, range, and range rate is detectedby a rendezvous radar system installed on tb e noseof tbe spacecraft . An orbit att i tude and maneuvering propulsion system (OAMS) util izing bypergol1cstorable propellants is installed in tbe adaptermodule and permits three axis attitude and trans lation control . The previously mentioned iner t ia lguidance system platform and computer units areutilized to convert tb e radar outputs into displayed thrust and att i tude commands wbicb enabletb e crew to accomplish tbe rendezvous maneuvers.The digi tal command system is used to receiveground commands to enable tb e crew to maneuver towitbin radar range of tb e target . Docking latchesmounted in the nose of th e spacecraft are util izedin conjunction with a docking adapter mounted onth e target vehicle to accomplish tb e f inal dockingoperation. Storage of the propellants and thrust ers again dictates us e of th e adapter as an equipment bay.

The fourtb mission objective, accomplisbmentof simplified countdown techniques, bas s ignif i cantly affected tb e design of tb e spacecraft . Anumber of tb e major SUbsystems sucb as tbe radar,re - entry att i tude control system, paraglider instal la t ion , fuel cel l and reactant system, cooling pumppackage, environmental control system, and maneuvering propellant system, have been bui l t into separatesubassemblies. This modular concept allows systemsto be cbecked ou t on tb e bench and quickly installedin the spacecraft. All of tbe elect r ical and elec tronics equipment in tb e re -entry module is in stalled in equipment bays easily ac cessible througbdoors in tbe outer mold line of tbe spacecraft.Test points ar e bui l t into a l l systems witb neces sary leads brougbt to conveniently accessibleconnectors fo r t ie - in to t es t equipment. Automaticcbeckout eqUipment is provided fo r rapid countdownoperation, an d a l l Aerospace Ground EqUipment hasbeen carefully integrated witb tb e spacecraft andlauncb pad systems .The f inal objective of establisbing man's pe r formance capabilities during extended per iods inorbi tal r l igbt bas le d to tbe basic concept of onboard mission command. Decisions and control capa bi l i ty are crew functions. The crew makes sucb de cisions as to when to abort, when to ini t iaterendezvous and retrograde maneuvers, with tb eground complex serving in a monitoring and advisorycapacity. Attitude and translation maneuvers aremanuallY controlled . All on -board systems ar emonitored and operated by tb e crew. The space

suits have been designed such tha t helmets, arms,and legs may be removed to appro ximate a shi r t sleeve operating condition. Provisions have beenmade in tb e hatches and th e pressurization systemto allow egress from th e cabin into space whenproperly suited.This concludes th e ro l l cal l of new featuresintroduced by th e specific Gemini mission object ives. To this l i s t can be added those items whichhave direct counterparts in the Mercury spacecraft.These include the basic aerodynamics shape and re entry heat protection concepts, th e l i fe support

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system utilizing a 5 psia oxygen atmosphere, UHFand HF voice communications, S- Band and C-Bandtracking beacons, solid propellant retrograderockets, re -entry module att i tude control thrustersystem, s i lver zinc batter ies fo r re-entry andpost - landing elect r ical power, and various recoveryaids.The remaining portion of tb i s paper will dis cuss th e integration of tb e foregoing features intotb e Gemini spacecraft design and wil l present amore detailed description of some of th e majorsystems.The ful l scale mockup photograph in Figure 1serves to relate tb e overall size of tb e spacecraftto tb e crewmen standing alongside. Worthy of notein tbis view are the inset individual windshieldsfo r th e pi lo t and crewman. In Figure 2, tb e mockup is arranged to i l lus t ra te tb e division pointsbetween tb e major structural assemblies . At th e

r ight of th e photograph is a 5- foot diameter targetdocking adapter which is supplied by McDonnell andis bolted to th e Lockheed Agena target vehicle.Next in l ine is th e re -entry module. I t consistsof a conical cabin section housing th e crew andmost of th e environmental control an d electronicseqUipment, surmounted by a cylindrical section con taining th e re -entry att i tude control thrusters towhich is attached th e rendezvous and re -entrysection in which the paraglider rendezvous radarand docking provisions are stowed. The sphericalsurface of th e ablative heat shield forms tb e baseof th e re -entry module. The overall length of th ere -entry module is 144 inches, and i t s maximumdiameter is 90 inches . For reference, corresponding Mercury dimensions are 90 inches and 74.5incbes, respectively. The adapter shown here intwo sections is actually bui l t as a single s tructural unit and severed during th e course of th emission into th e two parts i l lust rated . The com plete adapter is 90 inches long and tapers from th e120- inch Titan I I diameter a t one end to the 90 - inchre - entry module base diameter at th e other. Thepar t adjacent to th e re -entry module, termed th eretrograde section, contains th e four solid propellant retrograde rocket motors. The equipmentsection houses th e fuel cel ls and reactants, OAMSpropellants, coolant circulating pumps, and miscellaneous electronics and instrumentation equipment.The eqUipment section is Jett isoned just prior toth e retrograde maneuver by severing th e structure atth e point shown in th e photograph by means of af lexible linear shaped charge. A similar shapedcharge a t tb e 120- inch diameter base of the adapteris used to disconnect th e spacecraft from tb e launchvebicle af ter insertion into orbit. Attachment oftb e adapter to th e re -entry module is by 3 steelstraps spaced about tbe peripbery of th e re - entrymodule. These s tr aps, along witb wiring and tubing,ar e cu t simultaneously by shaped charges when th eretrograde section is jett isoned.

Figure 3 shows th e in ter ior arrangement of tb espacecraft . The crewmen s i t side by side but witheach seat canted outboard 12. This eliminates an ychance of contact during simultaneous seat ejectionand also conserves space fo r eqUipment . Thepressurized cabin area houses th e crew and tbei r


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The afterbody heat protection used on Geminiis almost identical to that proven on th e Mercuryspacecraft. As shown in Figure 7, high temperatureRene 41 shingles .016" thick are used over th econical section. Withstanding temperatures of upto 1800F, these shingles achieve a thermal balanceby radiation to th e atmosphere. The shingles ar eattached to th e basic structure using bolts andwashers through oversized holes to allow fo r therma l expansion. Small blocks of Min -K insulation areused a t th e support points, and a layer of Thermoflex insulation is used between supports to keepsubstructural temperatures within l imits . Over th ecylindrical sections of th e afterbody, heating ratesar e too high fo r efficient radiation cooling and,therefore, a heat sink principle is utilized. Beryllium shingles .24 inches thick on th e windward sid eand .09 inches thick on th e leeward side are in stal led over this area. Again, provisions fo rthermal expansion are included.

As a matter of in teres t , the heat distributionpattern over th e afterbody, based upon wind tunnelmodel data at a Mach number of 10 and an angle ofattack of 20, is shown in Figure 8 . The isothermsshown represent constant values of th e ratio of th elocal to the stagnation heat transfer coefficients,where

_ qstag- T -To wql = local heat transfer rate

stagnation point heat transfer rateTO free stream stagnation temperatureTw wall temperature at point considered

The more cri t ical conditions on th e cylindricalsection are apparent .The adapter structure i l lustrated in Figure 9consists of a cylindrical shell of HK -31 magnesiumskin .032" thick, stiffened by longitudinalstr ingers of HM-31 magnesium with stabilizing alu minum rings at several locations . As previouslymentioned, th e only other structural elements inth e basic adapter are the retrorocket support beamsshown in the Figure. Magnesium is util ized as th ebasic structural material in order to withstand

launch temperatures of up to 600F without furtherprotect ion. A unique feature of th e adapter, shownin th e sectional view, is th e manner in which th eentire outer surface is used as a space radiator.The environmental control system coolant is circu la ted through . 25" tubes whieh are extruded inte grally with th e longitudinal structural str ingers .Fifty foot long extrusions are doubled back andforth to form redundant coolant loops with a mini mum of connections . This arrangement not onlysaves weight, bu t resul ts in a superior design fromth e meteoroid puncture standpoint, since th e coolan t tubes are protected both by th e outer skin andth e legs of the extrusions .

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Environmental Control System. As noted inFigure 10 , th e basic concepts of the Gemini environmental control system are similar to those ofMercury. Points of similarity include th e us e ofa 5 psia pure oxYgen atmosphere, use of a spacesui t to back up th e pressurized cabin, CO2 removalby lithium hydroxide. Two significant departuresfrom th e Mercury system are incorporated in Gemini.These are the use of cryogenic rather than gaseousstorage fo r primary oxYgen, and th e us e of a coolant fluid and space radiator as th e primary meansof heat removal rather than water boiling.With th e exception of the radiator, th e en vironmental control system is supplied by th e

AiResearch Division of the Garrett Corporation .AiResearch also supplied th e Mercury system, and isthus able to draw upon this back- lo g of experience.The sui t and cabin are pressurized with oxYgensupplied from either th e primary cryogenic sourceor a secondary gaseous supply. The secondarysupply is stored at 5,000 psi in two 7 -lb . capacitybottles in th e re -entry module, and serves both asan emergency supply in orbit and as a normal supplyduring re -entry. Either of the two bott les wil lpermit a t least one fu l l orbit plus re-entry. Thecryogenic supply, stored in the adapter, containsup to 104 Ib s . of super -cr i t ica l oxYgen.The sui t compressors circulate th e oxYgenthrough an odor and CO2 adsorber, a heat exchanger,a water absorber, th e pressure suits, and a solidstrap . In th e event th e primary compressor becomesincapable of maintaining th e required circulationrate, a second redundant compressor is activated.A cabin fan circulates th e cabin atmosphere througha second heat exchanger.A silicon ester coolant, Monsanto MCS 198, isCirculated through th e cabin and sui t heat ex changers, eqUipment cold plates , fuel cel ls , andf inally th e space radiator to remove the heat ab sorbed. The coolant loop is completely redundant,and t wo pumps are provided in each loop . Normally,only one pump in one loop is required; however, fo rpeak elect r ical and solar load conditions, twopumps in one lo op , or one pump in each loop, areturned on . During launch , ae rodynam ic heat:fngraises th e tempe rature of th e adapter surface toth e point where th e radiatvr is ineffective.Therefore, i t is by -passed and th e coolant is cir culated through a water boiler during this phase ofth e mission . Approximately 30 minutes are requiredaf ter launch f or th e radiator to cool back down.The total water requirement for th is operation isless than 10 Ib s .The cabin is equipped with a dump valve toeffect depressurization, and a high flow rate repressurization valve. These, coupled with th esingle point hatch unlatching mechanism, providefor egress experiments in space . NASA is cu rr entlydeveloping th e necessary portable l i fe support k itfo r this application .Electrical System. A block diagram of th eelect r ical system is shown in Figure 11. Primaryelect r ical power during launch an d orbit phases ofth e mission is supplied from a hydrogen-oxYgen fuel

cel l battery stowed in th e adapter. This unit,

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currently under development by th e General ElectricCompany, is of th e ion exchange membrane type .Actually, th e installation in Gemini consists oftwo separate, identical packages - or sections, asthey are called - each of which ba s redundantcoolant loops, i ts own reactant control valves,electr ica l controls, and instrumentation. Eachsection is made up of three stacks of 32 individualcel ls . Each stack bas a rated output of 350 wattsa t 23 .3 volts fo r a to ta l rated power of 2100watts. No load voltage is 28 volts. I t is possible to shut down an y single stack in th e event amalfunction is detected. Peak power reqUirementsfor presently planned missions can be met by th efuel cel l battery even with one stack inoperative.

The cryogenic hydrogen and oxygen reactantstorage an d regulation system is supplied by th eAiResearch Division of the Garrett Corporation. Asin the case of th e breathing oxygen, two sizes oftanks ar e being developed with usage depending uponth e mission length.For retrograde, re - entry, and post-landingpbases of th e mission - which occur subsequent tojett isoning of th e equipment adapter - power issupplied from a bank of four 16 -cel l silver zincbatter ies rated a t 40 amp hours each . These bat

teries are t ied into the same main bus as th e fuelcel ls , an d serve as an emergency orbita l powersupply in case of a fuel cel l failure . In th eevent of a par t ia l fuel cel l failure, th e s i lverzinc batter ies may be used to augment th e fuelcel ls during th e few hours when peak power is re quired. This will permit successful completion ofth e mission even i f one complete section is los t .In th e event of complete fuel cel l system failure,th e batter ies wil l provide fo r at leas t one orbitfollowed by a normal re-entry an d a minimum of 12hours post-landing eqUipment operation.

A second battery system consisting of three15 ampere - hour 16-cel l silver zinc batter ies isprovided in th e re - entry module to power pyrotechniC devices and various control relays andsolenoids . Isolat ion of these systems from th emain bu s prevents feedback of voltage spikes,resulting from such devices, into cr i t ica l elec tronics equipment. This design resul ts fromMercury experience where such "gli tches" proved tobe a troublesome nuisance. As shown in th e blockdiagram, diodes are used to isolate th e two pyrotechniC squib batter ies from each other so tha tcomplete redundancy in pyrotechniC systems iscarried a l l th e way back to th e power source .

Another deviation from Mercury practice isthe provision of individual inverters for each ofth e several AC powered devices such as th e controlsystem electronics, iner t ia l guidance system, sui tand cabin fans, and coolant pumps. This allowselect r ical characteristics of each inverter to bematched in i ts particular applicat ion. Off designoperation with resulting penalties in conversionefficiency is thereby minimized.

Attitude and Maneuver Propulsion Systems. Ato ta l of 32 bi -propellant l iquid rocket thrustchambers ar e used fo r control l ing at t i tude andmaneuvering th e Gemini spacecraft . Thrust chambersizes an d locations are shown in Figure 12 . Asshown in th e l e f t -hand sketch, three independentat t i tude control systems ar e provided. Each con s is ts of eight 25 -lb . thrust units arranged to f i rein paral lel pairs for yaw an d pitch control or,different ia l ly , fo r ro l l . Two of th e systems arepackaged in the cylindrical re -entry control systemmodule a t th e forward end of th e cabin section.Each of these systems ba s i ts own propellant andpressurization tankage, valves, and l ines. Thesere - entry control systems, referred to as the RCS,are util ized only during th e retrograde an d re entry port ions of th e mission . They ar e maderedundant since they are considered essential tocrew safety.

The third r ing of at t i tude control thrustersis used during th e orbita l port ion of th e missionan d is located a t t h ~ rear of the adapter module .The eight maneuvering thrusters ar e arranged asshown in th e r ight - band sketch . Four lOO - lb .units are directed through th e center of gravityto provide fo r la tera l and vertical impulses. Apair of a ft -facing 100-lb . thrusters at th e baseof th e eqUipment adapter section provides forwardimpulse. A pair of 85 -lb . units facing forwardand canted slightly outboard, mounted on eitherside of th e adapter close to th e re - entry moduleattachment station, provides reverse thrust .

The adapter -mounted orbit a t t i tude and maneuve r propulsion systems, generally referred to asth e OAMS, share a common propellant supply. TheOAMS thruster arrangement permits attitude an d ma neuver control in the event of th e loss of an ysingle thruster . Complete redundancy is no t pr o vided because mission safety is no t directlyinvolved, and because of th e high weight penaltiesrequired to make a truly redundant system. Asindicated in th e typical section view in Figure 12,th e thrust chambers are ablatively cooled withceramic inserts a t th e throat section. Separatevalves are provided for fuel and oxidizer.

Propellants fo r both th e RCS and OAMS systemsare nitrogen tetroxide (N204 ) oxidizer and mono -methyl hydrazine ( N 2 ~ C ~ ) fuel, an d ar e pressurefe d to th e thrust chambers from bladder type tanks.Propellant capacity is 35 Ibs. for each of the twoRCS systems. This is sufficient to accomplishretrograde and re -entry with either system. Maxi mum propellant capacity of th e OAMS tanks isapproximately 700 Ib s . This is sufficient to provide at t i tude control throughout a rendezvousmission plus a 700 foot / second maneuvering velOCityincrement. For th e non - rendezvous, long durationmission, smaller OAMS tanks are used to save weightan d allow extra oxygen an d fuel cel l reactants tobe carried .

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The RCS and OAMS thrust chambers and propel lant systems ar e being supplied to McDonnell byth e Rocketdyne Division of North American Aviation .Guidance and Control Electronics. A detaileddescription of th e Gemini guidance and controlsystem and i ts operation is beyond th e scope ofth is report. A series of reports could be , and infact ha s been, written on this phase of th e space craf t design . Therefore, only a br ief descriptionof th e system and i ts components, along vi th a verycursory reviev of i ts functions, vi l l be attempted .Figure 14 presents a very simplified blockdiagram of th e guidance and control system. Pi lotinputs are made through either th e at t i tude controlor the maneuver control handle. These are proc essed through th e att i tude control and maneuverelectronics (ACME) vhich contains the logic c ir cuitry needed to select th e proper thruster valves .

Resulting spacecraft dynamics ar e sensed by rategyros, by th e horizon scanners and iner t ia l measuring unit of th e iner t ial guidance system and, inthe case of motion vi th respect to the target, byth e rendezvous radar. Outputs from th e sensingunits ar e fe d either directly or through th e com puter to the several displays to command pi lotaction. Depending upon th e particular control modeselected, outputs from th e sensors may also be fe ddirectly into the ACME to provide automatic ormixed control modes . In th e case of re -entry a t t i tude control, a direct mode exists in which controlsignals may be fe d directly from the control handleinto the thruster chamber solenoid valves.Responsibility fo r th e overall concept , defi nition and f inal integration of the guidance an d

control system into th e spacecraft rests vithMcDonnell. Component suppliers include MinneapolisHoneywell, Minneapolis, fo r th e ACME and rate gyrosystems ; Minneapolis - Honeywell, St . Petersburg, forth e inert ial measuring unit ; International BusinessMachine Corporation fo r th e computer and 6V indi cator ; Advanced Technology Laboratories for thehorizon sensors ; Westinghouse fo r the radar andrange rate display ; Lear -Siegler fu r th e attitudeand rate indicator ; an d Rocketdyne for the thrustersystem. IBM also ha s responsibil i ty for integra tion of th e overall inert ial guidance system and fo ranalYtical studies in areas of mission planningassociated vith computer programing .

Although no t shovn on th e diagram , a number ofredundancies exist in the system . For example,crew selectable backup units ar e provided fo r th ehorizon scanner, rate gyros, and attitude controlelectronics. As previously discussed, redundantre - entry att i tude control thrusters ar e availableand redundant control signals ar e provided to themfrom th e control handle .The pi lot ha s available three different manualattitude control modes: rate command, single pulse,and direct; and tv o automatic modes : orbi tal andre -entry . The rate command mode provides a space craf t angular rate which is proportional to th econtrol handle deflection. This mode is th e pri mary att i tude control mode used during maneuvers .With th e control handle centered in this mode,

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angular rates about a l l three axes are damped toless than .1 pe r second. Since .1 pe r second isequal to th e angular rate of the large hand on aclock, this mode is equivalent - fo r short timeperiods, at least - to an attitude hold mode. Thesingle pulse mode is one in vhich a single minimumduration thruster pulse results each time th e con t ro l handle is deflected from neutral. I t is usedfo r precise at t i tude control; fo r example, vhenpreparing to align th e iner t ial platform, or fo rmaking minor adjustments to angular rates duringextended orbital f l ight . The direct or f ly -by vire mode, as previously noted, is essentially abackup method of operating th e thrusters and re sults in a constant angular acceleration any timethe control handle is deflected .

The automatic orbital attitude control mode isa coarse slaving of the spacecraft att i tude to th evert ica l as detected by th e horizon scanner. I tholds rol l and pitch attitude to within approxi mately 5 during extended periods of orbi tal f l ightwithout th e need fo r operation of the completeinert ial guidance and associated electronics sys tems. Since no yaw reference is available in thismode , yaw at t i tude is manually controlled by thepi lot using th e single pulse mode fo r control andvisual observation of th e ground as a reference.The automatic re - entry at t i tude control mode pro vides rate damping in pitch an d yaw about th e aero dynamic trim point of th e spacecraft, an d rol latt i tude control in response to error signals fromth e inert ial guidance system.

Only one mode of translation thruster controlis provided . This is a manual direct control modein vhich thrusters ar e simply turned on or off bymotion of th e maneuver handle .The heart of th e guidance and control systemis th e iner t ial guidance system. As noted ear l ierin this report, i t is basically required to performth e mission objectives of rendezvous and controlledre -entry . However, as shovn in Figure 15 , i tsversat i l i ty ha s been exploited to perform a numberof other important mission functions. A most sig nificant one is backup fo r the basic radio guidanceused for th e Titan I I launch vehicle. Although no tindicated on th e block diagram, error signals fromth e l i lert ial guidance system may be fed to thelaunch vehicle autopilot to control i t s enginegimbal actuators . Another function to be pro gramed in th e computer is a launch abort naviga tion mode vhich vi l l enable touchdovn from launchabort to be made at pr e -selected points . Orbitalnavigation vhich is essentially keeping track ofpresent position in orbit may also be accomplished.This program is carried a step further to allovcalculation of th e retrograde time required totouch dovn at any pre - selected point vithin th emaneuver capabili ty of the spacecraft in the eventof an abort from orbit . Rendezvous maneuver com mands are generated vi th th e aid of th e rendezvoustracking and ranging radar. The re -entry controlprogram generates th e error signals necessary toaccomplish rol l att i tude control during re -entry.

Docking System. One of th e major objectives


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of th e Gemini spacecraft program is to actuallydock with th e target vehicle. The docking andlatching procedure is shown in Figure 16 . A ta rgetdocking adapter (TDA) supplied by McDonnell boltson to th e upper equipment bay of th e Agena targetvehicle. This adapter contains a radar transponderwhich operates in conjunction with the spacecraftrendezvous radar, flashing lights fo r visual targetacquisit ion, and a docking cone. The l a t te r is afunnel - shaped assembly supported by shock absorberswhich damp ou t impact loads and prevent rebound. AV-shaped s lo t in th e cone mates with an indexingbar on th e spacecraft to align th e two vehicles.This permits three latches in th e cone to engagef i t t ings in the nose of th e spacecraft. The coneis then retracted and locked t ight to rigidize th econnection between th e Gemini an d th e Agena. Theprocess is reversed to separate the two vehicles.The latching f i t t ings on th e spacecraft may beblown free by pyrotechnic charges as a backup meansof separation.The Gemini crew can command th e Agena att i tudeand propulsion systems as well as th e dockingmechanism by an RF link either before or af ter docking . Just prior to f inal docking and while attachedto th e Agena, the status of th e target vehicle isascertained from displays mounted above th e dockingring (not shown in the sketch).Landing System. The paraglider is th e onlymajor Gemini system being provided as governmentfurnished equipment to McDonnell. I t is beingdeveloped under a separate contract to NASA-MSC byNorth American's Space and Missile Systems Division .Detailed design insofar as installation in th espacecraft is concerned is being closely coordinatedwith that of the spacecraft. Certain portions ofth e system such as the gas supply fo r inflation ar ebeing provided by McDonnell.Figure 17 presents a br ief summary of th e para glider deployment and performance characteristics .The paraglider in conjunction with th e re -entrytrajectory control system permits touchdown a t th epre - selected si te . I t is flown much l ike a two control airplane using th e same hand controller asfo r orbit att i tude control.The landing gear, as previously described, isa tricycle skid type . The nose gear is extended a tparaglider deployment and th e main gear by pilotaction. The paraglider is jett isoned immediatelyaf ter touchdOwn to avoid possible interferenceduring th e run-out .In addition to th e paraglider, an 84- foot ringsa i l parachute recovery system is being developedby Northrop-Ventura under contract to McDonnell.This parachute will be util ized on th e unmannedlaunches, an d possibly several of th e early mannedf l ights, pending f inal qualif ication of the paraglider. Touchdown wil l be in water when th e para chute is used.Escape Provisions. Probably th e most noticeable difference between Mercury an d Gemini whenviewed externally is th e absence of t he escape tower

on th e l a t te r . In th e early part of this paper, th eaddition of th e ejection seat was l is ted as a designfeature added to back up the paraglider. Oncehaving accepted this requirement, extension of th eseat ' s capabilities to enable creW escape on th e padand during th e early phases of launch was natural.This was accomplished by incorporating a rocket - typecatapult in the seat. This gives sufficient a l t i tude and velocity to deploy th e parachute and landth e man a t a minimum of 600 feet from th e base ofth e launch vehicle. Application of th e ejection tooff - th e -pad operation is really made po ss ible by th efact that with th e Titan I I launch vehicle, there isno violent detonation associated with deflagrationof the propellants. Therefore, th e primary consideration is to achieve sufficient clearance fromth e f ire which might result from a launch vehiclemalfunction, rather than blast effects.

During the early phases of the launch, ejectionconditions ar e essentially th e same as from an a ir craft . In fact, th e maximum dynamic pressure ofapproximately 750 Ib s / sq . foot encountered during aGemini launch is only about one -half of th e valuewhich can be achieved with cu rr en t fighter typeaircraf t . However, as th e launch vehicle acceler -ates , temperature rather than pressure becomes th elimiting condition fo r ejection . A maximum a l t i tude of 70,000 feet has therefore been establishedfo r operation of th e Gemini ejection seat. As shownin Figure 18, there are two additional modes ofescape during th e launch phase. The f i r s t of these,which extends from th e ejection seat alt i tude of70,000 feet to an alt i tude of 522,000 feet, ut i l izesthe spacecraft retrorockets to separate the spacecraf t from th e launch vehicle. To achieve thiscapabili ty, th e ratio of maximum retrograde thrustto mass fo r th e Gemini spacecraft is increased toalmost double tha t of Mercury. In th e abort mode,th e 4 rockets are fired in salvo to give a to ta l of10,000 Ibs. of thrust. Above an alt i tude of 522,000feet, which corresponds to the point where th e veloc ity is approximately 20,000 feet / second, separationis accomplished in the same manner as a normal in jection in orbit. In this third mode , th e retro ~ o c k e t s are retained for possible use in th e normalretrograde mode to enable more f lexibil i ty in selec t ion of a touchdown si te .

In conclUSion, i t should be noted that spaceha s no t permitted a description of several im portant Gemini systems, including communications,retrorockets, and tracking systems . These are a l limportant, and a l l incorporate interesting features,but in concept have no t changed as drastically fromMercury as those covered in this report. However,future reports will most certainly deal in moredetai l with a l l Gemini systems.

7


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- - ~ - - - - - - -

MISSIONOBJECTIVES

NewDesignFeaturesResulting

8

TABLE 1

GEMINI SPACECRAFr

Long MeI1Il Life Offset Center of Orbital Attitude Modular ConceptGravityFuel Cells Radar AccessibleInertial Measuring EquipmentCryogenics Unit Iner t ial MeasuringUnit .Butl - in TestSpace Radiator Computer PointsComputerPCM Telemetry Digital COJIJlll6nd AutomaticSystem Digital Command CheckoutEquipment Adapter SystemParaglider Integrated AGEDocking

Landing Gear MechanismEjection Seats EquipmentAdapter

OTIIER FEATURES GENERALLY CORRESPOND TO MERCURY_

Figure 1. - Mock-up of Gemini spacecraft.

On-Boar dDecisionsManual ControlPart ial lyRemovableSpac e SuitEgress to Space


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Figure 2. - Division points between major structural assembly points of Gemini spacecraft.

/.I

///

I ,-RETROGRADE SECTION, ADAPTER

ADAPTER

Figure 3. - Gemini adapter and re-entry module.

9


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10

EL ECTRONICS RETRO GRADE SYSTEM

PARAGLIDER STOWAGE AND RA DAR

J- .)) OR liT-ATTITUDE AND MANEU VERSYSTEM TANKAGE\

RADAR QRYOGENIC IREATHI NG OXYGENENVI RONMENTAL CONTROL SYSTEMFigure 4. - Subassembly concept.

HEAT SHIELD

fun CELLS

SIDE EQUIPMENTACCESS DOORS

. / e n ~ . . - ' Ie--r-r/ E . ACCESS DOOR

FWD.IOTlOM EQUIP 1ACCISS DOOR 9? -- 'C. LANDING SKID DOOR

Figure 5. - Re-entry module structure .


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FIBERGLASS -

SIDEVIEW

TITANIUM

FIBERGLASSHONEy.cOMB

TITANIUM

FIBERITEMX 2625

FIBERGLASS HONEYCOMB FILLEDWITH DC-325 SILICON ELASTOMER

Figure 6. - Ablation shield.

0 000

FIBERGLASSCHANNEl

FIBERGLASS

~ ~ ~ b r INSULATION

INNER PANETEMPERED GLASS J ALUMINUM

WINDOW FRAMET Y ~ I C A L DOOR SPLICE

\ (\. -_.------IBERGLASS

.09 TO .24 BERYlliUM SHINGLES 7Figure 7. - Afterbody heat protection.

11


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12

203040so. 0

70

80

10 0

11 0

12 0uo140no16 0110

RE-ENTRY CONFIGURATION ___ 0'. =20BASED ON AEDC TUNNEL C DATA --- M .,. = 0Re "" 0 = .08 x 10

h LOCAL- - LINES OF CONSTANT HEATING = h STAG .

I I I I I I I I ! I I! I100 110 120 130 140 150 160 170 180 190 200 210 220 230

SPACECRAFT Z STATION -INCHESFigure 8. - Gemini afterbody heating distribution.

TYPICAL SECTIONA-A.2 5 IN. INSIDE DIA. /"' M-31 MAGNESIUM STRINGER

4 INCHES (APPROXI

Figure 9. - Adapter structure.


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-~ -

. - - -BASIC SYSTEM CONCEPTS SIMII.AR TO MERCURY:

CABIN AND SUIT ENVIRONMENT -PURE OXYGEN CABIN PRESSURE-S.l PSIA SUIT PRESSURE-3.5 PSIA EMERGENCY

3 IN . H2 0 BELOW CABIN NORMAL ORBIT OXYGEN-104 LBS. SUPERCRITICAL CRYOGENIC RE-ENTRY AND SECONDARY OXYGEN-DUAL GASEOUS

SUPPLY, 7 LBS. EACH SUITS ARE PARALLELED IN SINGLE CLOSED LOOP

CIRCULATING SYSTEM WATER REMOVAL BY WICK ABSORPTION OF WATER AT

THE HEAT EXCHANGER. COl REMOVAL BY LITHIUM HYDROXIDE BED

F igure 10 . - Envi r onmental control system.PRIMARY SYSTEM

HYDROGEN POTABLE& h r WATEROXYGEN I I '----___ 'PRE-LAUNCHAND ORBIT

SUPPLY

RE-ENTRY, POSTLANDING ANDEMERGENCYSUPPLY

ISOLATED SYSTEM

1 IDUAL FUEL/------

CELLBATTERIES

FOUR -ILVERI- -ZINC -BAnERIES

I SYSTEMMAINBUS ELECTRICALLOADS

DUAL DUALTHREE DUAL I--I- - - f ---- - I - ARMING _ Y R . Q T E C H N !SILVER ZINC SQUIB BUSES CONTROLS I-- CIRCUITS/ ------ - -BAnERIES

Figure 11CONTROL CONTROL

BUS I-- RELAYS &- Elect r ic al ower s stem. SOLENOIDS. p y13


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14

....

ATTITUDE CONTROL MANEUVER CONTROL100 LBS. THRUST PER UNIT5 LBS. THRUST PER UNIT

RE-ENTRYMODULE

(EXCEPT 8S lBS. FOR FORWARD FIRING UNITS)

ADAPTERMODUlE

Figure 12. - Thrust chamber arrangement.CONSTRUCTION OF " CRES " UNLESS OTHERWISE NOTED

+

RE-ENTRYMODULE

ORIENTED REFRASIL MOUNTING FLANGE& PHENOLIC

THROAT INSERT RANDOM REFRASIL(REFRACTORV) & PHENOLIC

NOTE:VALVE SECTION ROTATED 90% WITH RESPECTTO CHAMIER FOR CLARITV.

Figure 13. - Typical thrust chamber 25 lb . OAMS


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( PILOT DISPLAYS r - - - - - - - - - - - - - - - - - - - - - - - - - - - - -(SENSING & COMPUTING SYSTEM :IIIIII I IRANGE AND I I II I IRANGE RATE REND RADAR IIINDICATOR I II IRATE GYROS J---- IATTITUDE I II IAND I I

RATE Rl I IHORIZ SENSORj- II IINDICATOR l- I III I INERTIAL j II III I MEASURING IINDICATOR I UNIT II I II I ICOMPUTER II- - - - - -' I II INERTIAL GUIDANCE SYSTEM Il ______________ - - - - - - IIII_ _ __ J

,-------------------------- - - - - - - - - - - - - - - - - - - - ,I II 1I II------------- ----- --, I: CONTROL I I:n TTITUDE

ATTITUDE L. ATTITUDE IH- HANDLEPILOT CONTROL AND t--o AND HSPACECRAFTi: MANEUVER MANEUVER MANEUVER I DYNAMICSELECTRONICS THRUSTERS :

BACK-UPLAUNCHGUIDANCE

: CONTROL IIHANDLE II CONTROL SYSTEM II I~ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ __ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ __ _ _ __ _ JFigure 14. - Gemini guidance and control system.

CONTROLLEDLANDINGOBJECTIVE

RENDEZVOUSAND DOCKINGOBJECTIVE

INERTIAL GUIDANCESYSTEM

IMU-COMPUTER-DCS

LAUNCHABORT

NAVIGATIONMANEUVER

COMMANDS

ORBITALNAVIGATION

ORBITALABORT

NAVIGATION

RE-ENTRYCONTROL

Figure 15. - Inertial guidance system mission application.15


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IAPPROACH

Figure 16. - Docking phase.

DROGUE DEPLOYMENT ................ 60 ,00 0 FT . ALTITUDESTART PARAGLIDER DEPLOYMENT. . . . . . . . SO ,OOO FT. ALTITUDEPARAGLIDER DEPLOYMENT COMPLETE . . . . .40,000 FT . ALTITUDEGLIDE RANGE FROM 42,000 FT . L/ D=3 .1. .. . . 19.2 NA.MI.- DOWN RANGE

18 .2 NA.MI.- UP RANGELANDING VELOCITIES AT TOUCHDOWN . . . . 95 FT.lSEC. HORIZONTAL

10 FT. lSEC. VERTICAL (MAX . )RUNOUT DISTANCE AFTER LANDING . . .. 33 0 FT. APPROX.PILOT FLOWN USING HAND CONTROLLER, ACTUATES GAS OPERATED CA8LE REElS .

Figure 17. - Landing configuration.INnOION ALT. 529 .00 0 ft ., ' : 3305[(5.

MODE 3+ - - - + ! : : ~ : ~ ~ I ~ N ( ~ ! I I ~ ~ S - - - - - 522 .000 n ., t " J09 Uc .--..... { [OU",,,,[N.ITITAN THtUn

nRMINATlOPIIOItTOSlP.R.TlON

TOUCHDOWN CONTaOl THltUuno ,:11 0AMS2 ) IllTROOUOI loekOSJ I li n MODULAtION

70 ,000 . 1 :I 94 nc .MODE 1 UICTION SlAT

. n . O G I l A O ( AOAPTUSlCTlON

t b ~ OROGUIOEPLOYMENT

PAIt"GLIOlRD[PLOYMlNTAU .

' : : ;: :,ADAPTlIlSlenON

35 ,00 0 fT.----1I --

Figure 18. - Launch escape modes.

gemini design features

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