Description of the SERT II Spacecraft and Mission

8/8/2019 Description of the SERT II Spacecraft and Mission

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1. Report No. 2. Government Accession No.

DESCRIPTION O F SERT II SPACECRAFT AND MISSION

17. Key Words (Suggested by AuthorM )SpacecraftElect ric propulsionSERT II

7. Author(s)Richard G. Goldman, Guy S. Gurski, and William H. Ha wers aat

18. Distribution StatementUnclassified - unlimited

9. Performing Organization Name and AddressLewis Research CenterNational Aeronautics and Space AdministrationCleveland, Ohio 44135

12. Sponsoring Agency Name and AddressNational Aeronautics and Space AdministrationWashington, D. C. 20546

3. Recipient's Catalog Wo.

5. Report Date6. Performing Organization Code

November 1970 '

8. Performing Organization Report No.E-5807

10. Work Unit No.704-00

11. Contract or Grant No.

13. Type of Report and Period CoveredTechni cal Memorandum

14. Sponsoring Agency Code

20. Security Classif. (of this page)I Unclassified19. Security Classif. (of this report)Unclassified 21. No. of Pages 22. Price"24 I $3.00'For sale by the Clearinghouse for Federal Scientific and Technical Information

Springfield, Virginia 22151


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DESCRIPTION OF SERT I1 S P A C C R A F T AND MISSby R i c h a r d G. Goldman, Guy S. Gurski, a n d W i l l i a m H. Ha we rs a a t

L e wi s R e s e a r c h C e n t e rS U M M A R Y

This report is a descript ion of the SERT II spacecraft andA mission. The missionpurpose and the orbi tal flight plan to accomplish this purpose will be discussed. TheSERT II objectives a r e to prove the reliability and endurance capability of an ion th ru st er ,develop and refine operational procedure s , asce rtai n operating characteristi cs in thespace environment, and validate ground te st r esu lts . A miss ion objective of 6 monthsof continuous operation of one thr ust er system is expected to be achieved by a flight pro -gram str essin g the use and application of reliable, space-proven components and desi&philosophies.

INTRQDUCTIONThe SERT 11project was undertaken to substantiat e the reliability and endurancecapability of an ion thrus ter in the space environment, t o demonstrate the compatibility

of a thr ust er system capable of space missions with its associate spacecraft systems , todevelop and refine operational procedures , to asc erta in operating characteristic s in thespa ce environment, and to validate ground tes ts. A minimum mission time of 6 monthsw a s chosen as a suitable compromise among the l ife expectancy of flight-proven hard-ware, a continuous sunlight orbit , and operating time on a single thrust er'l ong enough tobe extrapolated to even longer missions such as would be required in powered flight toLthe other planets (refs. 1 o 3).

The SERT II satellite was launched from Vandenberg Air For ce Base on Februa ry 3 ,1970. This r eport describes the requirements of the mission, the orbit chosen to meetthe se requirements, and the space craf t in general, including experiment s, power system,tele metr y, command, and attitude control.


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Establishment of the charac teri stic operating life of the th rust er is the primary objective of the SERT I1 mission because it is character istic long operating life that transform s the ion thruste r from a laboratory device to a useful space thrus ter system. Thtest objective of 6 months was selected as a compromise between future mission requi rments and the cost, complexity, and organization available.

Also of significant importance is the validation of ground test ing provided by a spaflight. The mor e confidence that can be generated in simplified ground testing, the momoney can be saved by obviating the necessity of endurance tes ts in the actua l space environment. This saving als o reflec ts i n the elimination of effort required in running iothr ust ers in lar ge space simulation chambers.

Potential use rs of ion thr ust ers must be in possession of t he specialized techniquenec essar y to develop and operate them. The ground work fo r these specialized techniqhas been established through the SERT I and SERT 11 flight programs (refs. 4 to 9).

Finally, the actual operating characteristics of the ion thruster, as affected by thespace environment, cannot be accurat ely determined except in space (refs . 10 and 11).

Mission RequirementsThe mission requirements followed from the decision that the smallest t hru ste r

which would adequately demonstrate the operating par amete rs and endurance char act eristics of thi s type of engine wa s a l-kilowatt unit. Taking into account the sol ar ar r a ydegradation and spacecraf t housekeeping functions, an initial, on-orbit power of 1500watts was required. To power the thr ust er continuously requ ires a solar array flown ia 6 -month constant sunlight orbit. Sun synchronous orbit s ar e essentially pola r orbits.Constant sunlight is achieved by selecting the orbit altitude and inclination such that theoblateness of the Ear th, that is , the belt of extr a mas s distribution in the equatorial region, pre ce sse s the orbit plane at approximately 1 pe r day, that is , the angular rate awhich the Earth moves around the Sun. An on-orbit altitude of 1000 kilometers was s elekted as the result of an optimization of gravity gradient torques, aerodynamic dra g,solar cell degradation, required constant sunlight, and launch vehicle limitations.

The use of flight-proven hardware was an important ground rule laid down for theSERT I1 mission. Table I lists components derived from previous missions. In manycase s, identical units we re purchased from the manufacturer to the original specificatiand drawings. In the assembly of the experimental and prototype spacecraft, someactual su rplus hardware was obtained from the project concerned.be se en that considerable usage was made of previous designs. Although thi s resu lted

'

From the list it can

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TABLE I. - LIST O F COMPONENTS HAVINGPREVIOUS FLIGHT HISTORY

Components used in SERT 11dc-ac inverterSwitching mode regulatorPhase sensitive demodulatorsBattery chargerBatteryCommand decoderCommand receiverTransmitterHybridDiplexerTape recorderSubcommutatorMulticoderFrequency division multiplexerTime code generatordc-dc converterBackup acquisition systemControl moment gyroscopesSolar arra yMiniature electrostatic accelerometerHorizon scanner

Flight vehicleAgenaLEMLEMAgenaMarinerISISPegasusPegasusPegasusAir Fo rce programBIOSBIOSBIOSBIOSAgenaSurveyor and Lunar OrbiterAgenaAir Forc e programSaturnAgena

FR-1

significant cost sav ings, the m ost important f acto r gained was confidence in the reli-ability of the hardwar e based on its previous flight history.it could be modified to support the SERT ll mission with the large required solar arra y.The SERT-Agena, together with the spacecraft support unit (SSU) nd spacecraft (SC) ,provides excellent moments of iner tia fo r gravity gradient control.sive means using ther mal coatings on the outer surf ace s of the SC and SSU.

The Thorad-Agena was the launch vehicle. An advantage of the Agena vehicle is that

A study indicated that the th erm al control sys tem could be obtained by entirely pas-

F I ght SequenceAs indicated in the previous section, a 1000-kilometer constant sunlight orbit wa s

desirabl e to meet mission objectives. The Thorad vehicle was, therefore, launched fromthe Western Te st Range, which is the most efficient site from which to achieve a polarorientation. The Thorad burn plus first Agena burn put the s atel lite into an elliptical

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P i t c h - dow n maneuv e rfo r g r av i t y g r ad i en t''Iar u Agena s ec ond bu r ndeployment \ o r o r b i t i n j e c t i o n

/ Agena coast \

T hor - pow er ed f l i gh t1000 k i l ome te rpo l a r o r b i t

C S -39648F i gu r e 1. - R epr es en ta t i on o f Se r t I1 f l i ght sequence.

tra nsf er trajec tory to 1000-kilometers, where a second burn of the Agena circ ula riz edthe orbit. Figure 1 is a diagram of the injection procedure. The Agena vehicle attitudecontrol system oriented the sate llite nose down with the thr us te rs pointed toward Earth

The ion thrus ters , as operated in orbit, are offset by 10' from the vert ical . Th isoffset r esul ts in enough tangential thr ust t o r ai se or lower the orbit altitude, dependingon th ru st er selected. This altitude change is approximately 100 kilometer s during the6-month period. The orbit rais ing, which directly modifies orbit period, provides adir ect measurement of integrated thrust. Real-time thrus t was measure d by a miniatuelectrostatic accelerometer.

SATELLITE DESC R I TION

GeneralThe SERT II configuration is shown in figure 2 . The satellite is 1.525 mete rs in

diameter and 7 .92 me te rs long and weighs 1435 kilograms. Six mete rs of solar arr ayextend outward on both si de s of the Agena aft end equipment rack . As depicted in fig-ure 2 , the ion th ru st er s a r e pointed Earthward with the main bulk of the satellit e beingthe empty Agena. Attached to the forward end of the Agena is the SSU. The SSU pro-4


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Figure 2. - SERT-I1 i n orbit.5


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vides powel. conditioning, switching, tele met ry, and command syste ms . The pri maryattitude pontr ol syst em components are also in the SSU.perime nt, together with all associated experiments.

Forward of the SSU is the space craft (SC), which houses the prim e ion thr ust er ex-The general weight breakdown fo r each majo r assembly follows:

Weight,I kgSpacecraft 2 82Spacecraft support unit 220Agena launch vehicle 740Solar a r ra y -93

1435

Spacecraft Con st ruct onIdentified by the two protruding thr us te rs , the SC provides support and protection fo

the vari ous experiments . A completely redundant pa ir of ion th ru st er s is carried, eachwith its own power conditioning unit. Each th ru st er is mounted on gimbals so that thethru st v ecto r can be adjusted on orbit through the cente r of ma ss of the satellite, therebminimizing disturbance torques fr om this source. Several associated experiments arecarried; a space probe to measu re spacecr aft potentials, motor -driven beam probeswhich are swept through the ion beam on command t o meas ure the beam potential, su r-face contamination expe riments (solar cell s held at 71' and -43' C) to determine con-taminating efflux from the operating thru ster , a radiofrequency interference experiment(RFI) to me asur e any noise generated by the ion thr ust er in selected radio bands (thesebands will be used by interplanetary and orbiting spacecra ft), and finally, a miniatureelectros tatic accel eromete r (MESA) which m eas ures the ver y s ma ll acceleration of thesate llit e caused by the thrus t of th e ion engine.

Additional components satisfy the bas ic re quir ements for power switching, con-ditioning, and instrumentation. Spacecraft attitude is measur ed by horizon scanners . Abackup acquisition sys tem (BACS) is provided f o r reacquisition should orientation be los

The bas ic shape of the spacecraft is a right cir cular cylinder 53.4 ce ntimete rs highand 149 cen time ters in diame ter. Figure 3 shows the configuration and general constrution technique. The component ac ce ss plates form the skin of the spacec raft. Thre e ofthe sk ins a r e aluminum rat he r than the ligh ter magnesium because of improved compat-ibility of aluminum with the 2 -93 the rmal coating used. Magnesium alloy is used for fiv

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Figure 3. - Spacecraft struc ture .

outer sk ins t o sav e weight. The power conditioning units are mounted on an aluminumradiator which form s an integra l part of the spacecraft structure. The spacecraf t com-ponent locations are shown in figu res 4 to 7 .

Spacecraft Support Unit ConstructionThe SSU structure is very simi lar in construction, siz e, and appearance to the space-

craft, with the power conditioning radi ator being the exception. Seven tr ay s are used inth e SSU for conventional component mounting, whereas the spacecraft requires structuraladaptation for supporting the large experiments. Ther mal control patterns a r e simil arfor both spacecra ft and SSU. All component acces s areas are covered with skins si mi larto that on the spacecraft. Th ree skins have a 2-93 paint and five have perforated alu-minum tape. The SSU receives power from the sola r ar ra y, and with its power syste m,regulates and controls the housekeeping functions. The thr ust er power from the ar ra ypa sse s through the SSU to the spacecr aft power conditioners. The SSU also provides thetelemet ry and command functions required . It houses in its center section a controlmoment gyro package which or ien ts the sate llite in one axis and provides damping forall axes.

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A T ,-Attitude c o n t r o l nozz les

: B -d ' -R e fl ec to r e r os i on - T h r us te r pow er c ond i t i one r sA - e x p e r i m e n t,-Radiat ing sur face

,-Beam probe 2 e l ec t r on i c sI l -Space probe ele ctro nic s ,-c c e l e romete r e l ec t r on i c sI

Power switching Iunit B-3-l I ' , -Power sw i tch ingf/ / I unit B-2,- Beam pro be 1 l ec t r on i c s - ad i o f r equenc y i n te r f e r enc e

- P o w e r s w i t c h i n g unit B-1

- Ac tua to r mo to r c apac it o r s

Sect ion B-B( uppe r t r ay s )

' - H o r i z on s ens o r pow erand e l ec t r on i c s CD-9832-31Sec t i on A - A( l ow er t r ay s )

F i g u r e 4 - Gener a l a r r a ngem en t o f s pac ec ra ft .

a


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fge5-B8os

aas

as

u

fge6-B2os

aas

as

u

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F i gu r e 7. - Bay 4 o f s pac ec r a f t and s pac ec r a f t s uppo r t un i t ,

Spacecraft support unit component layout is shown in figures 5 to 8.

Launch VehicleThe Agena is modified to serve as a spacecra ft support platform. It has large ma s

distribu tions which provide favorabl e moments of i ner tia fo r gravity gradient stabiliza-tion. The SERT 11Agena is 6.2 5 mete rs long by 1. 525 mete rs in diameter, with a mod-

, ified aft equipment rack which supports and thermally protects the s ola r ar ra y.spacecraft Earthward, (2) deploying the large sola r a rr ay , (3) placing the SSU batteryon line, (4 ) dumping of all propellants aboard the vehicle , (5) maintaining proper attitudduring thi s sequence, and (6 ) (upon ground command) switching horizon scan ne rs tospacecraft control for attitude determination.

Once orbit is achieved, the Agena perf orm s se ve ral key functions: (1 ) pointing the

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Il V e h i c l ei n te r f ac e

Sw i t c h i ng mode r egu l a to r s - / - S i gna l c ond i t i one r

,-Time code ge ne rato r

M u l t i c o d e r s

-A- S u b c o m m u t a t o r s

C ommand r ec e i v e r s-.Command decoder --

Spacecraf ti n t e r f a c e

, - - T e l eme t r y c a l i b r a to r/- Vol tage-cont ro l led-os c i l l a to r m i x e rPow er c on t r o le l ec t r on i c s unit -

-- Te I m e t ry beaconCD-9379-31Lower t rays

Sec t i on A-A

F i g u r e 8. - Gener a l a r r ange me n t of s pac ecr a ft s up po r t unit.

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Pos i t ion A

01 1 ax i s// -Pitch axis: \\

/ w = Z r r a d l o r b i t// Yaw axis/////

/

\

\\\\ \\ \\\

III

IiS //\ /:\ /-/4 - -Polar c i r cu lar/

CD-8486-31

F i g u r e 9. - S E R T- I1 c i r c l i n g E a rt h.

Attitude ControlThe orientation of the satellite, depicted in figure 9 , is maintained in pitch and r olby gravity gradient and in yaw by gravity augmented by the control moment gyroscopes

(CMG). The sa te ll ite weighs 1435 kilograms and has a ma ss distribution s imil ar to abar bell, that is, idealized moments of i ner tia which r esul t in an optimum gravity gra-dient orientat ion capability. Moments of iner ti a are as follows:

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Resultant gravity gradient re stor ing torques for small angles are

The se r es tor ing torques maintain the spac ecra ft orientation stably in conjunction with theCMG's , which provide additional sti ffness in the yaw axis and the nec essary damping inall three axes.

A backup attitude control syst em (BACS) s provided in case a temporary disturbanceshould cause the satellite to lose orientation. Thi s BACS is a cold gas system which canreorient the spacecraft from disturbance ra te s up to 5' pe r second. Reacquisition is ac-complished with an open loop, by commands timed and executed by ground control.

Thermal SystemThe SERT 11 satellite uses a passive the rmal design because the satellite orientation

with re spect t o the Sun and Ea rth is relatively constant. Extensive analysis , modeltests, full scale tank tests, and experiments with coatings we re employed in the thermalcontrol design. This allowed the use of thermal coatings only (2-93, lack paint, pol-ished aluminum , and perforat ed aluminum tape). The the rma l control patte rn limits,sol ar and Earth t herm al inputs and regulates the radiated output so that a satisfactorythermal environment is maintained fo r all components. The th rust er power conditioners( P C l s ) are mounted on a large radiator which se rve s as part of the spacecraft structure.The therma l dissipation of th e radiat or requi res tha t , when both PC's a r e off, radiatorstrip heaters be turned on to maintain an adequate the rmal environment.

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F i gu r e 10. - Stowed s o l a r a r r ay .Power System

Solar array . - Rated at 1500 watts of usable power, the so lar a r r a y is divided intotwo functional supplies, a 60-volt t hrus te r section (1300 W) and a 35-volt housekeepingsection (200 W ). In the stowed position the a rr ay is attached to the rectangular tr us swork aft of the Agena tank section within the envelope of the Agena-Thorad adaptor. Figur e 10 shows the stowed array.

Deployment of the array is accomplished by springs and a sc iss or mechanism. Thea r r a y is fixed and so oriented on the orbita l vehicle tha t, when deployed, it lies withinthe orbit al plane facing the Sun. Figure 2 shows the a rr ay in the deployed position. Thear ray has a length of 6 me te rs on each si de of the Agena and a width of 1.525 me te rs . Itprovides an active area of 17.5 squa re me ter s.on the back of the a r ra y and reflective coating on the cover gla ss dn the f ront of th ear ra y, which maintain the ar ra y temp erat ure between 40' and 60' C depending on its

The ar ra y employs a modular construction, the basic module being 36.2 by 4 5 . 7

Therm al control on the ar ra y is achieved by the u se of tinted white acr yli c lacquer

, orientation.centimeters. Each module is composed of a thin magnesium waffle pla te to which arebonded 74 so la r cell submodules. The submodules are composed of five individual 2 - by2-centimeter sol ar cells connected in para llel on a 2- by 10-centimeter metall ic plate.The sol ar cells are 0. 3-millimeter-thick N/P cel ls with a 2-ohm-centimeter base re-sistivity and 11percent efficiency at air ma ss zero . Radiation and ther mal protectionis provided by 0. 5-millimeter-thick quartz cover glass. Six of the basi c modules a r e14


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incorporated into a panel which is 0.76 by I . 52 mete rs. The overall ar ra y consists of90 modules. The panel str uct ure to which th e modules are attached is lightweight alum-inum.

Power distribution. - The 60-volt section of the sol ar a rr ay provides power throughmotor driven switches to the ion th ru st er power conditioning syst em. The 35-volt sec -tion of the ar ra y is fed into switching mode regulato rs (SMR's) in the SSU for regulationand distribution. A block diagram of the power sys tem is shown in figur e 11.power to the spacecraft loads, inverters, battery charger, telemetry system, and signalconditioning. The standby SMR supplies power to the command sys tem and tr ansmit te rs .Diode connection of the main SMR to the standby SMR provides power t o the commandsystem and tran smit ters i f the standby SMR fails. Command capability exists to useboth o r either of the two SMR's to supply the spac ecra ft and SSU loads o r t o bypass bothSMR's and operate directly from the so lar a rr ay .

Two SMR's provide regulated 26.5 volts dc to *1 percent. The main SMR supplies

a r r a y(60 V,

Igenam lI

Spac ec r a f t s uppo r t unit

ar ray , (35 V,200W)

Aa t te r y br a n s m i t t e r s

Spacecraf tpow er c on -I

PowerHw i t c h i n gT h r u s t e rpow er c on -d i t o ne r

F i g u r e 11 - SERT I1 power system.

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Redundant 115 -volts ac, 400 -cycle inv ert er s operating from the main SMR, outputprovide power to any two of the fou r contro l moment gyroscopes in the SSU, o the beamprobe ac tuato rs and thr ust er gimbal motor s in the spac ecraft, and to thesensors.cles is electric ally back biased fro m the housekeeping sol ar a rr ay and is instantaneouslyavai lable fo r emergency conditions. Should the SSU sense an undervoltage condition, thebatt ery would come on line to support essenti al loads excluding the thr us te rs and experi-ments. A battery charger is provided to maintain charge status and rec harge the batteryshould it be required.

Power switching for t he te lemetry and power systems is provided in the SSU whileexperiment switching is performed in the spacecraft. The command rec eiv ers and de-coder cannot be switched off. All components in the spacecraf t and SSU are fused.Undervoltage protection is provided to di sable all sys tem s except the command systemand tr an smit te r should the SMR regulated output drop below 23 volts dc for longer than200 mill iseconds. The ion th ru st er power conditioning would be disconnected fro m thethr ust er ar r ay should the housekeeping arr ay dr op below 23 volts for longer than 1 sec-ond.

A 40-ampere-hour si lver oxide - zinc batt ery capable of at least five discharge cy-

Com municat on SystemTelemetry. - The air borne tele metr y sys tem, together with the supporting ground

station and communication network, fo rm s a composite system which provides theLewis control center with rea l-time telemet ry and command verification data duringperiods of ground coverage. In addition, data sto ra ge capability via two on-board tapere co rd er s in provided f or per iods when STADAN coverage is not available.

A simplified block diagram of the ai rbo rne telemetry system is shown in figure 12.Data from four subcommutators a r e fed into a multicoder, where the analog to digitalconversion and time division multiplexing are accomplished. The Dulse code modulated(PCM) ulticoder output is then fed into a voltage-controlled oscillator (VCO) through amixer and transmitte d at 136 megahertz as a phase modulated (PM) signal (PCM/FM/PM).command verification are also provided.tape reco rde r, and tran smit ter. Although the four subcommutators are not redundant,som e key data are on two subcommutators and a ls o data are allocated so that all datafrom one device ar e not on one subcommutator.

Timing pulse s for experiments and transmission of tape r eco rde r playbacks andRedundancy in the telemetr y syst em is provided for the multicoder , VCO'S, mixer,

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17.5 pulseslsect o co m m a n dsubsystem

ve r i f y s t a t u s -8. 75 Bi tslseccommand, ver i fy

To commandr e c e i v e r 14 i p lexer IVol tage con-t r o l l e d o s c i l -l a t o r s 1 an d

T i m i n g- l i n e s t o Command- xper i -m e n t s

lapi t s l sec I_Vol tage con-t r o l l e d o s c i l -l a t o r s 3 a n d4. IRIG 10Com m andPr i me sampled data f rom Tape re -spacecra f t suppor t unit c o r d e r s C o m m a n dand spacecra f t B-n d 2

M i x e r s1 n d 2

T r a n s m i t t e r 2I_)4 Dip lexerTo command '-+r e ce i ve r 2 270" DelayF i g u re 12. - A i rb o rn e t e l e m e t r y su bsys te m . ( I n t e r ra n g e in s t r u m e n t a t i o n g ro u p , IRIG.

Command system . - The airborne portion of the command system functions in con-junction with the track ing network to achieve real-tim e command control from the Lewisdata control center.spacecraft interfaces. Real-time commands are sent and executed by the operator atthe ground station in contact with the spacecraft.command capability . Upon receip t of a command, the r eceiv er aboard the spacecra ftdemodulates the radiofrequency signal and dir ect s the command signa l to the decoder,where the command received is identified, put in storage, and inserted in the telem etryoutput fo r transm issio n to the ground station. At the Lewis control center, the commandis verified as the cor rec t command and the STADAN operator is instructed to send anexecute signal to the sp acecraft.

Redundancy is provided in the command system by use of tw o command receiversoperated in parallel and a command decoder consisting of two fully redundant decoderPPha lves", each capable of decoding 108 commands. All essenti al commands fo r theth rust er experi ment and key housekeeping sys tem s are redundant.

Figure 13 is a block flow diag ram showing pr im ar y ground and

The command system is an amplitude modulated 148-megahertz syste m with a 216

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Command, veri fy, andrea l - t ime da ta d i sp lay

216 Comm

STADAN command,300 to 3000 Hzd u p l ex t e r m i n a l

andsaSERT 11 spacecra f t suppor t unit1 R e d u n d a n t co m m a n d d e cod e r

ICoup1 n gnetworks

R e d u n d a n txecute,TArX Command, Address, - co m m a n dre ce i ve rse lay sto rage recog n i t i ond r i v e r s

l 3 ; ? ' } M H i ,/r e q u e n c y s h i f t136.920 keyinglAMlAMtf

t r a n s m i t t e rre g i s t e rPu lse code

I IPulse co de m o d u l a t e d l f r e q u e n cy sh i f t ke y i n gF i g u re 13. -C o m m a n d sys te m f l o w d i ag ra m .

Antenna system. - A cir cularly polarized turnstil e configuration of four monopoleantennas with associated diplex ers and hybrids is used as the common telemetry commandantenna syste m.

ExperimentsThruster . - The ge neral configuration of the me rcu ry bombardment ion th ru st er is

shown in figur e 14. The th ru ste r assembly includes the propellant stor age , feed syste m,and neutra lizer (ref. 8). Permanent magnets are employed to create the longitudinalmagnetic fields. The feed system is a positive pre ss ur e type using gas behind a dia-phragm to force the mercur y from the res ervo ir. Both the thruster cathode and theneutra lizer a r e plasma dis charge device s utilizing hollow cathodes.18


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Gro

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The thruster requires 1000 w a t t s of power and produces 6. 2 millipounds of th rustwhen ope rated at the rat ed 250 mil liam pere ion beam curr ent and 3000 volts dc net ac-celerating potential. The design operating point of the th ru st er is 89 percent power effi-ciency and 76 percent propellan t utilizat ion, including neutr ali zer, at a specific impulseof 4240 seconds. Each of the two ion th ru st er sy st em s weighs 9 kilograms and carries15 kilograms of merc ury propellant.from the thr ust er portion of the s ola r ar ra y and conditions it to meet the requirementsof the thr us ter . Power requirem ents of the thr ust er are maintained ove r a solar arra yinput voltage range of 54 to 75 volt s. Average efficiency is 87 percen t at 60 volts inputwith the thrus ter operating at 250-milliampere beam current.construction as opposed to herme tic sealing. Unencapsulated magnetic components areused, elimina ting the need fo r void-free encapsulation. Each of the tw o power condition-ers weighs 14 .8 kilograms and is 26.6 by 51 by 1 7. 8 centimeters. The operating tem -perature range is 0' to 49' C.

The all solid state transis tor type inver ters operate at 8 kilohertz and provide thenine power supply requ irem ents of the th ru st er including the ne ces sar y control loops andtelemet ry outputs. A neu tral izer bias supply which is one of the thru st er expe riments isal so provided. Each of the individual power suppli es ha s the capability of withstandingsho rts between any power conditioner output te rmina ls and operating in a short circuitedcondition. Overload shutdown and input voltage undervoltage shutdown are als o provided.

Beam probe experiment. - It is required that the thru ste r beam plasm a potentialwith respe ct to the spacec raft be measured . Because of the high efflux of the plas maaccelera ted from the th rus ter and the consequent wear on the probe filament, a sweepingtype emissiv e probe is required (ref. 5). The probe m omentarily samp les the beam po-tential and then ret urns t o a position out of the beam.

When the beam probe is commanded to sweep, the probe electronics a r e automati-cally turned on to provide the ne cessa ry s igna l conditioning. The data ar e transmittedto the ground station in real time.

Each thrus ter has its own beam probe and actuator mounted on the thruster gimbalring assembly.

Space probe experiment. - In ord er t o evaluate the spacecraft potential during ionthru ster operation, a space probe (ref. 5) is required to determine the s pacecraft po-tential with respect to ambient space plas ma potential. It must be measur ed at a pointunaffected by the spacecra ft plas ma sheath and is, ther efor e, mounted on an isolated1.525 me te r boom, The boom is designed to deploy into the space plasm a on the leadingedge of the spacecraf t. Th is separati on se rv es to minimize spacec raft effects.

Surface contamination experiment. - The object of the s urfa ce contamination experi-

Power conditioning. - The power conditioner (r ef. 7) takes el ectr ical energy directly

The packaging configuration utilized f or the power conditioning is open-to-vacuum

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ment (ref. 6) is to determine if contamination exis ts in the vicinity of operating th rust er s.This experiment was formulated to determine the compatibility of so la r c ells to productsgenerated by ele ctri c thr ust ers .

Each surf ace contamination experiment (one for each thruster ) consists of two cellasse mbli es, one hot and one cold. Both a r e thermally isola ted from the struc ture andlocated near the ion beam.The solar cells a r e located dir ectly facing the Sun and produce th eir own signal(voltage), which is proportional t o the intensity of sunlight transmitted through the con-taminant coating on the sur fac e of the cell assembli es .

Miniature electros tatic acceleromet er experiment. - The object of this experimentis to obtain a n accurat e measu re of the thrus t impart ed to the spacecraf t during themission. There are four modes of ion thr ust er operation of part icular interest: thr ust eroff, 30 percent beam, 80 percent beam, and 100 percent beam. The acce leromete r outputper mit s an accu rate determination of thes e levels.

The accelerometer (ref. 9) permit s instantaneous measurement of the thrust in realtime. This measurement is provided from the acc eleromet er electronics in both analog(coarse ) and digital (accurate) form.noise generated by a n ion bombardment thruster system has resulted in the RFI experi-ment. The experiment receive s predetermined frequency bands and establishes thepower levels at these frequencies. A wideband antenna which views the ion beam ofeither thruster is required to receive the se signals . Frequency bands, which have beenselected, reflect spacecraft communication syste ms design for application in future deepspace missions. The frequency bands are 300 to 700 megaher tz, 1680 to 1720 mega-hert z, and 2090 to 2130 megahertz .

Reflector erosion experiment. - The reflector erosion experiment determines thedegradation of a reflecto r surf ace in nea r Ear th orbit. The optical degradation causedby the impact of micrometeoroids is measured as a change, during the course of themission , of the temper atu re of a polished plate. Since the disk is oriented toward theSun, with a relatively constant thermal input, the temperature shift of the disk isolatedfrom the spacecraft is a direct indication of opt ical degradation.

Radiofrequency interference experiment. - The need for identifying the type of RF

CONCLUDING REMARKSThe SERT II program w a s required to provide a long life test of an elec tric thruster ,

obtain operational experience in this new method of propulsion, and measure the charac-teri stic s of e lectric thrusters as they affected the design of future electrically propelledspacecraft.

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The satellite consists of a spacecr aft support unit in which a r e concentrated thetelemetry and command functions, ac and dc power conversion equipment, and mostswitching and housekeeping functions as well as control moment gyros. The principalexperiments, two ion thrust ers, each with its own specia lized power conditioning, to -gether with subsidiary experiments to closely monitor the char acte rist ics of the thr ust -e r s , are mounted in the s pace craf t, which bolts to the spacecraft support unit. The sub-sidiary experiments c ompris e beam probes and a spac e probe, radiofrequency inte r-ference, miniature electrostatic accelero meter, contamination cells, and reflector e ro -sion unit.attitude control and supports the s olar ar ra y and spacecraft - spacecra ft support unitcombination. The resulting satellite represe nts a simple, low-cost solution to the basi crequirem ents laid down fo r the SERT 11mission. To date, the spacecraft has performedsatisfact orily and prelimi nary data indicate that all mission objectives will be m et.

The Agena launch vehicle provides prop er m oment s of i ner tia for gravity gradient

Lewis Research Center,National Aeronautics and Space Administration,

Cleveland, Ohio, July 13, 1970,704-00.

REFERENCES1. Stuhlinger, Er nst : Ion Propulsion fo r Space Flight. McGraw-Hill Book Co ., Inc .,

1964.2. Zola, Charles L. : Interplanetary Probe Missions with So lar -Electric Propulsion

System s. NASA TN D-5293, 1969.3. Mullin, J. P. ; Barber, T. ; and Zola, C. : A Survey of Solar Powered Ele ctri c Pr o-

pulsion fo r Automated Missions. Pa pe r 68-1120, AIM, Oct. 1968.4. Rulis, Raymond J . : Design Considerations and Requirements fo r Integrating an

Elec tric Propulsion System into the SERT II and Future Spacecraft . NASA TMX-52853, 1970. (Also paper pr esented at the 8th AIAA Elec. Prop. Conf., Stanford,Calif., Aug. 31 - Sept. 2 , 1970.)

5. Jones, Sanford G. ; Staskus, John V. ; and Byers, David C. : Preliminary Results ofSERT 11Spacec raft Potential Measurements Using Hot Wir e Emi ssive Probes. NASATM X-52856, 1970. (Also paper presented at the 8th AIAA Elec. Prop. Conf.,Stanford, Ca lif ., Aug. 31 - Sept. 2 , 1970.)

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6. Staskus, John V. ; and Burns, Robert J. : Depositions of Ion Thr us te r Effluents onSERT 11Spacecraft Sur faces. NASA TM X-52860, 1970. (Also pa pe r presented atthe 8th AIAA Elec. Prop. Conf. , Stanford, Calif. , Aug. 31 - Sept. 2, 1970.)

7. Bagwell, J am es W. ; Hoffman, Anthony C. ; Leser, Robert J. ; Reader, Karl F. ;Stover, John B. ; and Vasicek, Ftichard W . : Review of SERT 11 Power Conditioning.NASA TM X-52858, 1970. (Also pa pe r presen ted at the 8th A I M Elec. Prop. Conf. ,Stanford, Cali f., Aug. 31 - Sept. 2, 1970.)

8. Kerslake, William R. ; Byers, David C. ; Rawlin, Vincent K. ; Jones, Sanford G. ;and Berkopec, Fran k D. : Flight and Ground Performance for the SERT 11Thrust-er. NASA TM X-52848, 1970. (Also pape r pr esented at the 8th A I M Elec. Prop.Conf. , Stanford, Ca lif ., Aug. 31 - Sept. 2 , 1970.)

9. Neiberding, William C. ; Lesco, David J . ; and Berkopec, Frank D . : ComparativeIn-Flight Thru st Measurement of the SERT 11Th ru st er . NASA TM X-52839, 1970.(Also paper pre sented at the 8th AIAA Elec. Prop. Conf. , Stanford, Calif. ,Aug. 31 - Sept. 2, 1970.)

10. Gold, Harold; Rulis, Raymond J. ; Maruna, Frank A. , Jr . ; and Hawersaat,William H. : Descrip tion and Opera tion of Spacecraf t in SERT I Ion Thrustor FlightTe st . NASA TM X-1077, 1965.

11. Cybulski, Ronald J . ; Shellhammer, Daniel M. ; Lovell, Robert R. ; Domino, EdwardJ. ; and Kotnik, Joseph T . : Results fr om SERT I Ion Rocket Flight Tes t. NASATN D-2718, 1965.

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Description of the SERT II Spacecraft and Mission

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