A Synopsis of Ion Propulsion Development Projects in the US SERT 1 to Deep Space I

8/8/2019 A Synopsis of Ion Propulsion Development Projects in the US SERT 1 to Deep Space I

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NASA/TM--1999-209439 AIAA 99-2270

A Synopsis of Ion Propulsion DevelopmentProjects in the United States:SERT I to Deep Space I

James S. Sovey, Vincent K. Rawlin, and Michael J. PattersonGlenn Research Center, Cleveland, Ohio

Prepared for the35th Joint Propulsion Conference and Exhibitcosponsored by the AIAA, ASME, SAE, and ASEELos Angeles, California, June 20-24, 1999

National Aeronautics andSpace Administration

Glenn Research Center

October 1999


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A Synopsis of Ion Propulsion Development Projects in the United States:SERT I to Deep Space

James S. Sovey, Vincent K. Rawlin. and Michael J. PattersonNational Aeronautics and Space Administration

Glenn Research CenterCleveland, Ohio 44135

Abstract GSFCHSHugheshe historical background and characteristics of

the experimental flights of ion propulsionsystems and the major ground-based technology HVdemonstrations were reviewed. The results of the IAPSfirst successful ion engine flight in 1964, SERTI which demonstrated ion beam neutralization, are IPSdiscussed along with the extended operation of JPLSERT II starting in 1970. These results together LeRCwith the technology employed on the early NSSKcesium engine flights, the ATS series, and the NSTARground-test demonstrations, have provided theevolutionary path for the development of xenonion thruster component technologies, control PASsystems, and power circuit implementations. In PPUthe 1997-1999 period, the communication rmssatellite flights using ion engine systems and the RPMDeep Space I flight confirmed that these SATMEXauxiliary and primary propulsion systems have S/Cadvanced to a high-level of flight-readiness. SCATHA

Acronyms and Abbreviations SEPSA communication satellitebuilt by Hughes for SESApplications TechnologySatelliteBreadboard Power ProcessingUnitDefense Advanced ResearchProjects AgencyDigital Control and InterfaceUnitDeep Space IEngineering ModelElectromagnetic InterferenceElectro Optical SystemsEast-West StationkeepingAcceleration due to Earth'sgravity, 9.807 m/s 2Glenn Research Center

ASTRA 2AATSBBPPUDARPADCIUDS1EMEM1EOSEWSKgGRC

SEPSTSESSERTSITSNAPSNAPSHOT

SPIBSST4USAFXIPS

Gc,ddard Space Flight CenterHughes SpacecraftHughes Space andCommunications CompanyHigh VoltageIon Auxiliary PropulsionSystemIon Propulsion SystemJet Propulsion LaboratoryLewis Research CenterNorth-South StationkeepingNASA Solar ElectricPropulsion TechnologyApplications ReadinessPanAmSat CorporationPower Processing Unitroot mean squareRevolutions Per MinuteSatmcx of Mexico CompanySpacecraftSpacecraft Charging at HighAltitudeSolar Electric PropulsionSystemSolar Electric PropulsionSystem TechnologySociete Europeenne desSatellites of LuxembourgSpace Electric Rocket TestStructurally Integrated ThrusterSystems for Nuclear AuxiliaryPowerSpacecraft carrying the SNAP10A nuclear power supply anda cesium ion propulsionsystemSatellite Positive-lon-BeamSystemSpace Technology 4United States Air ForceXenon Ion Propulsion System

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Introduction

Kilowatt-class ion propulsion systems havefound applications tbr spacecraft North-Southstationkeeping, orbit insertion, and primarypropulsion for deep space missions._'2 The firstsuccessful ion propulsion flight system wasdemonstrated in 1964 during the course of theSERT I flight. 3 Later on seven more ionpropulsion systems and one ion source systemwere flown as space experiments using mercury,cesium, or xenon expellants. A total of sixsuccessful, operational flights of IPS are now inuse for NSSK or primary propulsion. Articleson the origins of ion propulsion can be found inReferences 4 and 5.

Surveys of the history of electric propulsionsystems have cataloged the evolution of IPStechnology and generally described many of theexperimental and operational flights. _ Thepurpose of this paper is to provide more detailrelated to the IPS flights and major grounddemonstrations of the technology. Back_oundon system perlbrmance and in-space operationwill be summarized, and the evolution ofelectron-bombardment ion thruster developmentin the United States will be discussed.

Experimental Fli_hts of IonPronulsions Systems

The experimental flights of ion propulsionsystems developed in the United States aresummarized. Some of the results indicated in theTables Ia and I b are expanded, and major resultsare described. Although there were maior groundtest and development programs associated witheach of the experimental flights, nearly all of thesynopsized results reported here are associated'with the end product which is the flight test.pro2ram 661A. Test Code A

In November of 1961, Electro-Optical Systemswas awarded a contract by the U. S. Air Force todevelop a 8.9 mN, cesium contact ion propulsionsystem ['or three sub-orbital flight tests. TheElectric Propulsion Space Tests were calledProgram 661A and were managed by the AirForce Space Systems Command in LosAngeles. l"-_-'

The cesium contact engine incorporated anionizer array of 84 porous tungsten buttons. Thepower level, thrust, and specific impulse were0.77 kW, 8.9 mN, and 7400 s, respectively inthis engine which had a beam extraction diameterof about 7 cm. The neutralizer was a wirefilament which was not immersed in the ionbeam. Power to the PPU was supplied by 56 Vbatteries. The longest ground test was 1230hours.The first sub-orbital flight test was launched onDecember 18, 1962. When the high-voltagepower supplies were first turned-on, intermittenthigh-voltage breakdowns occurred, and the beampower supply became inoperative. Post-flightexamination of the power supply indicated thehigh-voltage breakdowns were probably causedby pressure buildup in the PPU due to gas ventedfrom the spacecraft batteries. The PPU high-voltage section was not adequately vented to keepthe pressure low enough. Engine thrusting wasnot accomplished in this test.SERTIThe SERT I spacecraft was launched July 20,1964 using a Scout vehicle. __3 This flightexperiment had a 8 cm diameter cesium contaction engine and a 10 cm mercury electronbombardment ion engine and was the firstsuccessful flight test of ion propulsion. Thecesium engine was designed to operate at 0.6 kWand provide 5.6 naN of thrust and a specificimpulse of 8050 s. The cesium flow wascontrolled by a boiler and the porous tungstenionizer electrode. The 10 cm diameter mercuryengine provided flow control via a boiler and aporous stainless steel plug. A hot tantalum wirewas used as the discharge cathode. Beam andaccelerator power supply voltages were 2500 Vand 2000 V, respectively. The engine had a 1.4kW power level with 28 mN of thrust at aspecific impulse of about 4900 s. Ion beamneutralization was provided by a heated tantalumfilament.

The early part of the flight was dedicated toattempts to operate the cesium engine. Thecesium engine could not be started because of ahigh-voltage electrical short circuit. The mercuryengine was started about 14 minutes into theflight. The IPS was successfully operated for

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31minuteswith53high-voltageecycleventswhichwerehandledythePPUfaultprotectionsystem.Eachoftherecycleventss onlya t;ewsecondsuration.Majorresultsromthetestwerehefirstdemonstrationf anIPSin space,effectiveonbeamneutralization,oEMIeffectson other spacecraftystems,and effectiverecoveryromelectricalreakdowns.hrustwasmeasuredsing threeindependenteasuringsystems,ndtherewereno majordifferencesbetweenn-spacendgroundestperformance.program 661A. Test Code BTest Code B was the second in the series of threesuborbital flight tests of the Electro OpticalSystems' 8.9 mN, cesium ion enginesystems. "'H A Scout vehicle launched thepayload on August 29, 1964. The launch wasdesigned to provide about 30 minutes above analtitude of 370 km. Alter 7 minutes into theflight, the engine was operated with ion beamextraction. Full beam current of 94 mA wasachieved about 10 minutes later. During thecourse of engine operation, an electric fieldstrength meter was used to infer payload floatingpotential relative to space. Spacecraft potentialwas about 1000 V negative during most of theengine operation with the filament neutralizer.The absolute value of payload potential wasabout ten times higher than anticipated, and it issuspected that there was inadequate neutralizationof the ion beam. The contact ion engine operatedlbr approximately 19 minutes until spacecraft re-entry into the atmosphere.In addition to withstanding the environmentalrigors of space flight, the IPS demonstratedelectromagnetic compatibility with otherspacecraft subsystems and the ability to regulateand control a desired thrust level.Program 661A. Test Code C

The third and final IPS payload of the Air Force'sProgram 661A was launched on December 2t,1964. __'* In this test, an additional wireneutralizer was incorporated and was immersed inthe ion beam to provide a higher probability ofadequate neutralization. The contact ion engineonly achieved about 20% of full-thrust before re-entry into the atmosphere. The short test timewas due to a very short burn of the Scout

vehicle's third stage. The high voltage wasapplied to the engine 7 minutes into the flightwhen the altitude was 490 km. Engine operationended after 4 minutes when the altitude was only80 kin.

SNAPSHOTOn April 3, 1965 a SNAP 10A nuclear powersystem was launched into a 1300 km orbit with acesium ion engine as a secondary payload. _5-_7The ion beam power supply was operated at4500 V and 80 mA to produce a thrust of about8.5 mN. The neutralizer was a barium-oxidecoated wire filament. The ion engine was to beoperated off batteries tbr about one hour, and thenthe batteries were to be charged lbr approximately15 hours using 0.1 kW of the nominal 0.5 kWSNAP system as the power supply. The SNAPpower system operated successfully for about43 days, but the ion engine operated for a periodof less than I hour betbre being commanded offpermanently. Analysis of flight data indicated asignificant number of high-voltage breakdownswhich apparently caused sufficient EMI to inducefalse horizon sensor signals which created severeattitude perturbations of the spacecraft. Groundtests indicated that the engine arcing producedconducted and radiated EMI significantly abovedesign levels. It was concluded that lowfrequency, < 1 MHz, conducted EMI caused theslewing of the spacecraft .aTS-4Two cesium contact ion engines were launchedaboard the ATS-4 spacecraft on August 10, 1968.Flight test objectives were to measure thrust andto examine electromagnetic compatibility withother spacecraft subsystems/'mr" The 5 cmdiameter thrusters were designed to operate at0.02 kW and provide about 89 _aN thrust at about6700 s specific impulse. Thrusters had thecapability to operate at 5 setpoints from 18 [aNto 89 l-tN. Thrusters were configured so theycould be used for East-West stationkeeping.Prior to launch, a 5 cm cesium thruster was lifetested for 2245 hours at the 67 laN thrust level}"During the launch process the Centaur stage didnot achieve a second burn. and the spacecraftremained attached to the Centaur in a 218 km by760 km orbit. It was estimated that the pressure

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atthesealtitudeswasbetweenx 106Torrand1x 10_Torr_. Eachof thetwoengineswastestedonat leastwooccasionsachoverthethrottlingange.Combinedesttimeof thetwoenginesasabout10hoursovera55dayperiod.The spacecrafte-enteredhe atmospherenOctober17,1968.TheATS-4lightwasthefirstsuccessfulrbitaltestofanionengine.TherewasnoevidencefIPS electromagneticnterferenceelatedtospacecraftubsystems.Measuredaluesofneutralizermissionurrentweremuchessthanthe ion beamcurrentimplying inadequateneutralization.The spacecraftotentialwasabout132V whichwasmuchdifferenthantheanticipatedalueofabout40V.IS

A flightIPS.denticalotheoneflownonATS-4. was launched on ATS-5 on August 12, 1969.The purpose of this flight was to demonstrateNSSK of a geosynchronous satellite. -_'-'-_ Oncein geosynchronous orbit the spacecraft could notbe despun as planned, and thus the spacecraftgravity gradient stabilization could not beimplemented. The spacecraft spin rate was about76 revolutions per minute, and this caused aneffective 4g acceleration on the cesium feedsystem. The high g-loading on the cesium feedsystem caused flooding of the discharge chamber,and normal operation of the thruster with ionbeam extraction could not be perlbrmed. TheIPS was able to be operated as a ncutral plasmasource, without high-voltage ion extraction,along with the wire neutralizer to examinespacecraft charging effects. The neutralizer wasalso operated by itself to provide electroninjection for the spacecraft charging experiments.SERT 11The SERT II development program which startedin 1966 included thruster ground tests of6742hours and 5169 hours duration. Aprototype version of the SERT II spacecraft wasground-tested tor a period of 2400 hours with anoperating ion engine. The spacecraft waslaunched into a 1000 km high polar orbit onFebruary 3. 1970. -'_ In addition to diagnosticequipment and related IPS hardware, thespacecraft had two identical 15 cm diameter,

mercury ion engines and two PPUs. Flightobjectives included in-space operation for a periodof 6 months, measurement of thrust, anddemonstration of electromagnetic compatibility.The thruster maximum power level was0.85 kW, and this provided operation at a 28 mNthrust level at 4200 s specific impulse. Flightdata were obtained from 1970 to 1981 with anion engine operating intermittently in one ofthree different modes, namely, HV ion extraction,discharge chamber operation only, or justneutralizer operation.Major results were that two mercury enginesthrusted for periods of 3781 hours ard2011 hours. Test duration was limited due toshorts in the ion optical system. Thrustmeasured in space and on the ground agreedwithin the measurement uncertainties. Up to300 thruster restarts were demonstrated. A PPUaccumulated nearly 17,9(X) hours during thecourse of the mission. Additionally, the IPS waselectromagnetically compatible with all otherspacecraft systems.ATS-6The purpose of the ATS-6 flight experiment wasto demonstrate NSSK of a geosynchronoussatellite using two cesium ion enginesystems, zL22242s Thruster development testsincluded a lifetest of 2614 hours and 471 cycles.Thruster input power was 0.15 kW whichresulted in a thrust of 4.5 mN at a specificimpulse of 2500 s. The ATS-6 was launched onMay 30. 1974. One of the ion engines operatedfor about one hour and the other for 92 hours.Both of the engines fail_ to provide thrust onthe restarts due to discharge chamber cesiumflooding. The feed system flooding problemcaused overloading of the discharge and highvoltage power supplies. This failure mechanismwas verified through a series of ground tests. 25

The IPS operation demonstrated an absence ofEMI related to spacecraft systems, verifiedpredictions of spacecraft potential with enginesoperating, and demonstrated compatibility withthe S/C star tracker. It was found that the ionengines or just the neutralizer could dischargelarge negative spacecraft potentials at all times.Further, tests indicated that "differential chargingwas reduced by the neutralizer when operated in

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spotmodeandeliminatedyoperationf theionengine.-SCATHA. P78-2

The SCATHA spacecraft had two charged particleinjection systems one of which was the SatellitePositive-Ion-Beam System (SPIBS). 2627 Thiswas a xenon ion source which included some ofthe technologies used in thrusters: however, thedischarge chamber was not performanceoptimized as was done with ion engines.Maximum operating power was 0.045 kW, andthe ion source could produce a thrust of about0.14 mN at a specific impulse of 350 s. Ionscould be ejected at 1 keV or 2 keV. Neutral-ization was accomplished by a tantalum filament.The specific impulse is low because there was noattempt to optimize the propellant efficiency.The SPIBS system was ground-tested for a periodof 600 hours. The SCATHA spacecraft waslaunched January 30, 1979 and placed in a neargeosynehronous orbit. Ion beam operations wereperformed intermittently over a 247 day period.The SCATHA flight demonstrated that "a chargedspacecraft, and the dielectric surfaces on it, couldbe safely discharged by emitting a very lowenergy (


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6.8kgofmercurywhichcouldprovideperationat full-powerbr approximately0,000hours.Theenvelopeasabout31cmlongby 12 cmdiameter. The SIT-5 development programfocused on the thruster and li_xt systemdevelopment: there was no PPU technologyeffort.Hollow cathode component tests demonstratedover 2800 simulated duty cycles. A separate testof the SIT-5 thruster was conducted for9715 hours at a beam voltage of 1300 V, athrust of 1.8 mN, and a specific impulse of2500 s. 3_3_ During the initial 2023 hours, thethruster was operated with a translating screengrid thrust vector system, and the thruster had anelectrostatic thrust vector system for theremainder of the test. The electrostatic beamvector grids were operated at 5 degrees deflectionfor about 120 hours and at either 2 degrees or4 degrees deflection for 1770 hours. There werean number of grid shorts that were successfullycleared by the application of 200 V to 400 V atcurrents from 6 mA to 70 mA. These tests werehelpful in the later definition of grid-clear circuitsfor the lAPS, XIPS, and NSTAR thrusters.

The SIT-5 mercury propellant system wassuccessfully tested lor a period of 5400 hours inan independent test.

SEPSThe Solar Electric Propulsion Stage program wasstarted in the early 1970s with a goal to provide aprimary ion propulsion system capable ofoperating at a fixed power for Earth orbitalapplications or over a wide power profile such aswould be encountered in planetary missions.One of the potential planetary targets was anencounter with the comet Enke. 36_7 The SEPSprogram included the development of 25 kWsolar arrays, PPUs. thermal control systems,gimbals, throttleable/long-life 30 cm diameterion thrusters, and mercury propellant storage anddistributions systems. This multi-Center, multi-Contractor effort was ongoing for about 10 yearswith a NASA investment of approximately $30million dollars. Because of funding limitations,a planetary flight program was not carried out:rather, a ground-based technology demonstrationwas pursued.

The thrust subsystem was a bi-module consistingof two thrusters, two PPUs, a propellant system,a gimbal system, thermal control, and supportingstructure. 3_39 This module would be a basicbuilding-block of a electric stage with simpleinterfaces. The 30 cm thruster was designed for2.6 kW input power with 128 mN thrust and aspecific impulse of about 3000 s. 5'3_ Thethruster/PPU was capable of throttling down to1.1 kW. More detailed references related to thedevelopment and test of the SEPS bi-modulehardware can be found in Reference 37.One of the early engineering model thrusters wastested for 10,000 hours over an input powerrange of 0.8 kW to 2.4 kW. a', Endurance tests ofthese 30 cm ion engines confirmed the need forspalling control of sputter-deposited dischargechamber coatings, 4"4_ and for the need of lowsputter-yield materials for the cladding of pole-pieces and baffles. 42 Other tests indicated thatvery small concentrations of nitrogen in thevacuum facility could significantly reduce wearon the upstream surface of the screen gridcompared to that expected in space. 4_Subsequent to these EM thruster tests, a total ofseven advanced engineering model thrusters weretested in segments including 3,940 hours and5,070 hours and a total test time of14,541 hours. _2 Ninety five percent of the testwas implemented using either breadboard orbrassboard PPUs which were of the series-resonant inverter design. _'_44

The Ion Auxiliary Propulsion System project andother preflight technology work took place in the1974 to 1983 timeframe. 45 Flight test objectiveswere to verify in space the thrust duration,cycling, and dual thruster operations required lbrstationkeeping, drag makeup, station change, andattitude control. This implied demonstration ofoverall thrusting times of 7,000 hours and 2500on/off cycles. The 8 cm diameter, mercury ionengine input power was 0.13 kW, and the thrustwas 5.1 mN at a specific impulse of 2500 s.The masses of the flight thruster-gimbal-beamshield unit, the PPU, and the digitalcontroller were 3.77 kg, 6.85 kg, and 4.31 kg,respectively. 4_' The system stored 8.63 kg of

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mercury,andthe propellanttorageandti,_:x:lsystemweighed.56kg. ThelAPSsuccessfullycompletedll flight qualificationestsandwasinstallednanAir Forceechnologyatellite.7Theflight of theTealRubyspacecraftascanceledytheAirForcedueolackof funding.Duringhecoursef thetechnologyndpreflightprogramsherewereanumberof enduranceestperformed. laboratory-typecmenginewastestedot 15,040hoursand460cyclesat the0.14kWlevel.4_ AnengineeringodelAPSengineandPPUweresuccessfullyestedor9,489hoursand652cycles.'_ThethrusterndPPU were located in the same vacuum chamberduring this test. In another test, an engineeringmodel thruster was operated at full-thrust for7112 hours and had 2571 restarts. 5_' There wereno major changes in thruster performance, and nolife-limiting degradation effects were observed. Asingle PPU was used to run two tests and hadoperated tor 14,000 hours without malfunction.XIPS-25 _1.3 kW_

This XIPS-25 program, conducted by HughesResearch Laboratories, developed thrusters,BBPPUs, and a feed system pressure regulator forpossible NSSK of 2500 kg class communicationsatellites. 5_ The 25 cm diameter, 3-grid, xenonion engine input power was 1.3 kW with a thrustlevel of 63 mN and a specific impulse of 28(X1 s.Three thruster types were developed, namely, alaboratory-type, an advanced development model,and an engineering model. Performance testsindicated that the later models inherited virtuallyidentical performance. A BBPPU with greatlyreduced parts count, over SEPS designs, wasbuilt and tested. Overall PPU efficiency was90c_, and the flight packaged specific mass wasestimated to be 8 kg/kW. A 15 month wear testwas conducted using the laboratory modelthruster, a BBPPU, and a flight-type regulator.The hardware successfully completed 4,350 hoursof testing and 3850 cycles which is equivalent toabout 10 years of NSSK. The Hughes Space andCommunications Company subsequentlypursued development of XIPS-13 (0.44 kW) andXIPS-25 (4.2 kW) systems, instead of the 1.3kW XIPS-25 system, for NSSK and orbitinsertion applications.

XIPS-25 (4.2 kW)

A 25 cm diameter xenon engine system is beingdeveloped by the Hughes Space andCommunications Company for NSSK, EWSK,attitude control, and momentum dumping for itsHS 702 spacecraft, s2-_'_The ion thrusters providestationkeeping at a cost of only 5 kg per year.Additionally, the IPS will be capable of boostingthe communication satellite's 14,500 km perigeeof the initial elliptical orbit to a circulargeosynchronous orbit. Chemical propellantsavings could be as much as 450 kg. It isplanned that the HS 702 spacecraft use fourXIPS-25 engines and two PPUs. Only two ofthe four thrusters are required to perform thestationkeeping and momentum control functions.Hughes has not yet launched the HS 702spacecraft with the XIPS-25. The spacecraft hasan end-of-life solar array power capability ofabout 15 kW.

Each thruster has an input power of 4.2 kW andprovides 165 mN thrust at 3800 s specificimpulse. XIPS hardware is currently underextended tests at the Hughes-Torrance. CA 6.1 mdiameter by 12.2 m long vacuum facility.Evolution of Electron-BombardmentIon Thruster Development in the

United States

Figure 1 shows a "roadmap" of the history ofelectron-bombardment ion thruster developmentin the United States from the first tests of a 10cm engine 55 to the first operational flights in1997/1998. TM Much of the history of the earlydevelopment of mercury ion engines was outlinedin Reference 5. The following remarks onlyfocus on the insertion of componentimprovements to the mercury and xenon ionengines. In the early 1960s the wire-grids werereplaced by multiaperture grids, s6 Later, in themid-1960s engine life extension was madepossible by the incorporation of hollow cathodesfor the neutralizer and main discharge, sT-s9 TheSERT II flight was the major in-spacedemonstration of these technologies. 2_ Majortechnology improvements in the 1970s were thedevelopment of high-perveance, dished grids 6_,methods to control spalling of sputter deposited

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materialn thedischargehamber1,andmethodsto providedeep-throttling.Mercuryenginesweredevelopedithdiametersangingrom5cmto 150cm. Enduranceestsof theseenginesextendedfrom about 4,000 hours to15,000hours.In the1980imeframet wasdecidedo replacethemercurynginesithxenonenginesecausexenonwaslesscontaminatingo spacecraftsurfaces,ndground-testperationseregreatlysimplified.In the 1980sand1990sing-cuspdischargehambers_'62wereusedinsteadofdivergent-fieldhambershosepole-pieces,nthevicinityof thedischargehamberathode,sufferedevereon erosion. Thering-cuspchamberso not requirepole-piecesn thevicinityof thehollowcathode,ndheboundarymagneticielddeviceeducesheion lossesothechamberwalls._ Additionally,ong-life,xenonhollowcathodeechnologyasenhancedbydevelopmentsn theSpaceStationPlasmaContactorrogramwhichfocusedn definingreliableprocessing,andlingandtestproceduresforthecathodes)Groundestsof 13cmand30cmdiameterenonenginesemonstratedorethan8,000hoursof reliableoperation,4_'shecommunication satellite and deep space tests ofthese engines, starting in 1997, confirmed thethrusters and PPUs are a very mature technology.Overational Flights of Ion Provulsion

Systems

In 1997/1998, a new era of ion propulsion forspacecraft began with the deployment ofcommunication satellites using an IPS with0.44 kW thrusters tbr auxiliary propulsion and adeep space mission using a 2.3 kW thruster.These were the first operational uses of IPS byUnited States industry and government.PAS-5. _;alaxv Vlll-i. ASTRA-2A.SATMEX 5. PAS B

As shown in Table 3, the Hughes Space andCommunications Company has launched fiveoperational communications satellites eachemploying four-0.44 kW xenon ion thrusters forNSSK. s-'s_ The high specific impulse IPSreduces the propellant requirements, versuschemical systems, by 300 kg to 400 kg, thusallowing incorporation of more communications

hardware aboard the spacecraft or reduction inlaunch size and cost. The IPS consists of twofully redundant strings each consisting of twothrusters and one PPU. Two daily "burns" of5 hours each are generally required for the NSSKfunction. Typical spacecraft lifetime is about15 years.Approximate masses for a thruster and PPU are5.0 kg and 6.8 kg, respectively. _ Overall IPSdry mass for the spacecraft is about 68 kg. ThePPU contains seven power modules for thebeam, accelerator, discharge, two keepersdischarges, and two heaters. Overall PPUefficiency of a BBPPU was 88%.PanAmSat was Hughes' first customer for theXIPS-13 propulsion system on PAS-5. Thiswas the first successful, operational spacecraftemploying IPS and was launched August 27,1997 from Kazakhstan on a Russian Protonrocket. Since then, four more spacecraft areoperational using the XIPS-13 system namely,Galaxy VllI-i, ASTRA-2A, SATMEX 5, andPAS 6B.Deer Svace 1

The NSTAR program provided a single string,primary IPS to the Deep Space I spacecraft. The30 cm ion thruster operates over a 0.5 kW to2.3 kW input power range providing thrust from19 mN to 92 mN. The specific impulse rangesfrom 1900 s at 0.5 kW to 3100 s at 2.3 kW.The flight thruster and PPU design requirementswere derived with .the aid of about 50development tests and a series of wear-tests atNASA LeRC and JPL of 2000 hours, 1000hours, and 8193 hours using engineering modelthrusters, 26s The flight-set masses for thethruster. PPU, and DCIU were 8.2 kg, 14.77 kg,and 2.51 kg, respectively 67. About 1.7 kg masswas added to the PPU top plate to satisfy theDSI micrometeoroid requirements. The powercable between the thruster and PPU wascomprised of two segments which were connectedat a field junction. The thruster cable mass was0.95 kg, and the PPU cable mass was 0.77 kg.The xenon storage and feed system dry mass wasabout 20.5 kg. A total of 82 kg of xenon wasloaded for the flight. Thrusters and PPUs weremanufactured by Hughes, and the DCIU wasbuilt by Spectrum Astro, Inc. The feed system

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developmentasa collaborativeffortbetweenJPLandMoog,nc/'sAsof April 27, 1999, the primary thrusting ofthe NSTAR engine system required to encounter

the asteroid 1992KD was completed. Thethrusting time at the end of April was 1764hours. Thruster input power levels were variedfrom 0.48 kWto 1.94 kW. Atotal of 11.6 kgof xenon was expended. As shown in Table 4,the NSTAR engine already has demonstrated thelargest propellant throughput in space ascompared to a SERT II engine that expendedabout 9 kg of mercury. Propellant throughput isa signature of total impulse capability. Nearly70 kg of xenon remains on the DS 1 spacecraftfor a possible mission extension. It is intendedthat the DS 1 spacecraft will pass within 10 kmof the asteroid 1992KD in July 1999. If anextended mission is approved, DS 1 willencounter comets Wilson-Harrington and Borreltyin the year 2001.

Next G_neration Ion PropulsionZ.t.thsl..el.q.xix_

Over the next decade, it is expected that there willbc many communications spacecraft employingthe XIPS-13 and XIPS-25 propulsion systems.Additionally, the Space Technology 4 spacecraftwill be developed by JPL for a flight to thecomet Tempel 1. and a small vehicle will be sentto the comet surface for scientific measurements.The ST4 spacecraft will use three NSTAR ionengine/PPU subsystems for primary propulsion.In the next few years, new IPS technologies willbe developed by NASA for higher thrust density30 cm ion engines and sub-kilowatt, smallerengines both of which have application toplanetary and Earth-orbital spacecraft. Some ofthe near-term work, shown in Figure 2, involvesdevelopment of titanium and carbon-carbon ionoptics which will provide significant lifetimeimprovements compared to the baselinemolybdenum grid systems. Low-power and Iow-flowrate neutralizers are also needed tor a wideclass of thrusters which operate at low powerlevels or are throttled over a wide range of inputpower. Design approaches and manufacturingtechnologies which provide reduced ion engineand PPU mass and cost are receiving significantattention in order to enable or enhance planetary

and small-body missions using relatively smalllaunch vehicles.

Conclusions

The historical background and characteristics ofthe experimental flights of ion propulsionsystems and the major ground-based technologydemonstrations were reviewed. The results of thefirst successful ion engine flight in 1964. SERTI which demonstrated ion beam neutralization, arediscussed along with the extended operation ofSERT II starting in 1970. These results togetherwith the technology employed on the earlycesium engine flights, the ATS series, and theground-test demonstrations, have provided theevolutionary path for the development of xenonion thruster component technologies, controlsystems, and power circuit implementations. Inthe 1997-1999 period, the communicationsatellite flights using ion engine systems and theDeep Space 1 flight confirmed that theseauxiliary and primary propulsion systems haveadvanced to a high-level of flight-readiness.

References

1. Beattie, J. R., Williams, J. D., and Robson,R. R., "Flight Qualification of an 18-raN Xenonlon Thruster," IEPC Paper 93-106, September1993.2. Sovey, J. S. et al., "Development of an IonThruster and Power Processor for NewMillennium's Deep Space 1 Mission," AIAAPaper 97-2778, July 1997.3. Cybulski, R. J., et al., "Results from SERT IIon Rocket Flight Test," NASA TN D-2818,March 1965.

4. Anon., "Ion Propulsion, Over 50 Years inthe Making,"http://science.nasa.gov/newhome/headlines/ion-ast.htm. February 1999.5. Anon., "A Case History of TechnologyTransfer." NASA TM-82618. August 1981.6. Holcomb, L. B., "Survey of SatelliteAuxiliary Electric Propulsion Systems," Journalof Spacecraft and Rockets. Vol. 9, No. 3, March1972.

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7. Molitor, J. H., "Ion PropulsionFlightExperience,Life Tests, and ReliabilityEstimates," Journal of Spacecraft and Rockets,Vol. II, No. 10, October 1974.8. Pollard, J. E., et al., "Electric Propulsion

Flight Experience and Technology Readiness,"AIAA Paper 93-2221, June 1993.9. Maninez-Sanchez, M. and Pollard, J. E.,"Spacecraft Electric Propulsion-An Overview,"Journal of Propulsion and Power, Vol. 14,No. 5, September-October 1998.10. Davis, J., "Sub-orbital Flight Testing ofElectric Propulsion Systems," proceedings of theSymposium of Advanced Propulsion Concepts.Science Publishers, Inc., January 1966,pp. 1-20.11. Tannen. P. D., "Engineering Support IorElectric Propulsion Space Tests," in AFSC llthAnnual Air" Force Science atut EngineeringSymposium, June 1964,12. Ernstene. M. P., et al., "Surface IonizationEngine Development." Journal of Spacecraft andRockets, Vol. 3. No. 5, May 1966,pp. 744-747.13. Gold, H., et al., "Description and Operationof Spacecraft in SERT I Ion Thruster FlightTest," NASA TMX- 1077, March 1965.14. Tannen, P. D. and Radoy, C. H., "ElectricPropulsion Space Tests," USAF TechnicalReport Number AFSWC TR-65-2, June 1965.15. Brunings, J. E. and Johnson, C. E.,"Nuclear Power in Space," Mechanical Engin-eering, February 1967, pp. 35-41.16. Davis, J. D. and Burnett. J. R., "RadiationHardening of an Ion Propulsion System," inRecord of the 1965 htternational Symposium onSpace Electronics. November 1965, pp. 13-B1to 13-B 16.17. Sellen. J. M., "Interaction of SpacecraftScience and Engineering Subsystems withElectric Propulsion Systems," AIAA Paper69- I i 06, October 1969.

18. Hunter, R. E., et al., " Cesium Contact IonMicrothruster Experiment aboard ApplicationsTechnology Satellite (ATS)-IV," Journal ofSpacecraft and Rockets, Vol. 6, No. 9,September 1969, pp. 968-970.

19. Worlock, R., et al., "An Advanced ContactIon Microthruster System," Journal of Spacecraftand Rockets, Vol. 6, No. 4, April 1969,pp. 424-429.20. James, E. L. and i3oldner, S. J., "Ion EngineSystems Testing," AFAPL-TR-69112, February1970.

21. Bartlett, R. O., et al., "Spacecraft ChargingControl Demonstration at GeosynchronousAltitude," AIAA Paper 75-359, March 1975.22. Olsen, R. C., "Experiments in ChargeControl at Geosynchronous Orbit--ATS-5 andATS-6," Journal of Spacecraft and RocketsVol. 22, No. 3, May-June 1985.23. Kerslake, W. R. and Ignaczak, L. R.,"Development and Flight History of SERT 1ISpacecraft," AIAA Paper 92-3516. July 1992.24. James, E. L., et al., "A North-SouthStationkeeping Ion Thruster System forATS-F," AIAA Paper 73-I133, October 1973.25. Worlock, R. M., et al., "The CesiumBombardment Engine North-South Station-keeping Experiment on ATS-6," AIAA Paper75-363, March 1975.26. Masek, T. D. and Cohen, H. A., "SatellitePositive-lon-Beam System," Journal of Space-craft and Rockets, Vol. 15, No. I.January-February 1978, pp. 27-33.27. Olsen, O. C., "Investigation of Beam-Plasma Interactions," Final Report, NASACR-180579, May 1987.28. Shuman, B. M. and Cohen, H. A.,"Automatic Charge Control System torSatellites," NASA Conference Publication 2359,also AFGL-TR-85-O018, papers from theSpacecraft Environmental InteractionsTechnology Conference held in October 1983.

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29.Masek,T. D. andPawlik,E. V., "ThrustSystem Technologylbr Solar ElectricPropulsion,"IAAPaper8-541,une1968.30.Macie,T.W.,etal.,"Integrationf aFlightPrototypeowerConditioneritha 20-cmIonThruster,"IAAPaper1-159,anuary971.31.Hyman, ., "DesignandDevelopmentf aSmall StructurallyIntegratedon ThrusterSystem,"ASACR-120821,ctober1971.32.Hyman,., "Pertormance Optimized, SmallStructurally Integrated Ion Thruster System,"NASA CR-121183, May 1973.33. Nakanishi, S., et al.. "Status of a Five-Centimeter-Diameter Ion Thruster TechnologyProgram," AIAA Paper 71-690, June 1971.34. Nakanishi, S., "Durability Tests of s Five-Centimeter Ion Thruster System," AIAA Paper72-1151, November 1972.35. Nakanishi, S. and Finke, R. C., "A9700-Hour Durability Test of a Five CentimeterDiameter Ion Thruster," AIAA Paper 73-1111,November 1973.

36. Duxbury, J. H., "A Solar-Electric Spacecraftfor the Encke Slow Flyby Mission," ALAAPaper 73-1126, November 1973.37. Anon., "30-Centimeter Ion ThrustSubsystem Design Manual," NASA TM-7919 I,June 1979.

38. Sharp, G. R., "Thruster Subsystem Moduletor Solar Electric Propulsion." Journal ofSpacecraft and Rockets, Vol. 13, No. 2. February1976, pp. 106-110.39. Schnelker, D. E. and Collett. C. R., "30-cmEngineering Model Thruster Design andQualification Tests," AIAA Paper 75-341,March 1975.40. Collett, C. R., et al.. "Thruster EnduranceTest," NASA CR-135011, May 1976.

41. Power, J. L. and Hiznay, D. J., "Solutionsfor Discharge Chamber Sputtering and AnodeDeposit Spalling in Small Mercury IonThrusters," AIAA Paper 75-399, March 1975.42. Bechtel, R. T., et al., "Results of theMission Profile Life Test," AIAA Paper82-1905, November 1982.43. Rawlin, V. K. and Mantenieks, M. A.,"Effect of Facility Background Gases on InternalErosion of the 30-cm HG Ion Thruster," AIAAPaper 78-665, April 1978.44. Bless, J. J., et al., "Electric Prototype PowerProcessor for a 30 cm Ion Thruster," NASACR-135287, March 1977.45. Power. J. L., "Planned Flight Test of aMercury Ion Auxiliary PropulsionSystem--Objectives, System Descriptions, andMission Operations," AIAA Paper 78-647, April1978.46. Collett. C. R., "Auxiliary PropulsionSystem Flight Package," NASA CR-180828,November 1987.47. Smith, B. A., "Teal Ruby Spacecraft to bePut in Storage at Norton AFB," Aviation Week& Space Technology, January 8, 1990,pp. 22-23.48. Nakanishi, S., "A 15,000-Hour CyclicEndurance Test of an 8-Centimeter-DiameterElectron Bombardment Mercury Ion Thruster,"NASA TMX-73508, November 1976.49. Dulgeroff, C. R., et al., "lAPS (8-cm) IonThruster Cyclic Endurance Test," IEPC Paper84-37, May 1984.50. Francisco, D. R.. et al., "SuccessfulCompletion of a Cyclic Ground Test of aMercury Ion Auxiliary Propulsion System,"IEPC Paper 88-035, October 1988.51. Beattie, J. R., et al., "Status of Xenon IonPropulsion Technology," A1AA Paper 87-1003,May 1987.

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52.Beattie,. R., "XIPSKeepsSatellitesonTrack,"The Industrial Physicist, June 1998,pp. 24-26.53. Anon., "Power to Burn: Versatile NewSeries Answers Customer Needs,"http://www.hughespace.com/factsheets/702/702.html.54. Anon.. "XIPS: The Latest Thrust inPropulsion Technology,"http://www.hughespace.com/factsheets/xips/xips.html.55. Kaufman, H. R., "An Ion Rocket with anElectron-Bombardment Ion Source," NASA TND-585, 1961.56. Kaufman, H. R. and Reader, P. D.,"Experimental Performance of Ion RocketsEmploying Electron-Bombardment Ion Sources,"in Progress in Astronautics and Rocketry,Voi. 5, Electrostatic Propulsion, AcademicPress. Inc., 1961, pp. 3-20.57. Sellen, J. M. and Kemp, R. F., "Researchon Ion Beam Diagnostics." NASA CR-54692,1966.

58. Sohl, G., Fosnight, V. V., and Goldner, S.J., "Electron Bombardment Cesium lon EngineSystem." NASA CR-5471 I, April 1967.59. Rawlin, V. K. and Pawlik, E. V., "AMercury Plasma-Bridge Neutralizer," Journal ofSpacecraft and Rockets, Vol. 5, February, 1968,pp. 159-164.60. Rawlin, V. K., Banks, B. A., and Byers, D.C., "Design, Fabrication, and Operation ofDished Accelerator Grids on a 30 cm IonThruster." AIAA Paper 72--486, 1972.

61. Moore, R. D., "Magneto-ElectrostaticallyContained Plasma Ion Thruster," AIAA Paper69-260, March 1969.

62. Sovey, J. S., "Improved Ion ContainmentUsing a Ring-Cusp Ion Thruster," Journal ofSpacecraft and Rockets, Vol. 21, No. 5,September-October 1984, pp. 488-495.63. Matossian, J. N. and Beattie, J. R.,"Characteristics of Ring-Cusp DischargeChambers," Journal of Propulsion and Power,Vol. 7, No. 6, November-December 1991,pp. 968-974.64. Patterson, M. J., et al., "Plasma ContactorTechnology for Space Station Freedom," AIAAPaper 93-2228, June 1993.65. Polk, J. E., et al., "The Effect of EngineWear on Performance in the NSTAR 8000 HourIon Engine Endurance Test," AIAA Paper97-3387, July 1997.66. Beattie, J. R., Williams, J. D., and Robson,R. R., "Flight Qualification of an 18-mN XenonIon Thruster," IEPC Paper 93-106, September1993.

67. Gronroos, H. G., NSTAR Project Office atJPL, Private Communication, May 1998.68. Bushway, E. D., Engelbrecht. C. S.. "andGanapathi, G. B., "NSTAR Ion Engine XenonFeed System: Introduction to System Design andDevelopment," IEPC Paper 97-044. August.1997.

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Table la. Experimental Flights of Ion Propulsion Systems

Spacecraft Program SERT I Program 661A, Program SNAPSHOT661A, Test Test Code B 661A, TestCode A Code C

S p o n s o r USAF NASA LeRC USAF USAF USAFBuilder of IPS EOS LeRC HuLzhes EOS EOS EOS

Launch date 12.18.62 07.20.64 08.29.64 12.21.64 04.03.65

Orbit, km Suborbital Suborbital Suborbital Suborbital

IPS type Contact Contact Contactionization ionization ionization

Pro p ella n t Cesium Cesium CesiumNo. of thrusters 1

Thruster anode -7 cmdiameter

Type of Wire fi lamentneutralizer

5000 Veam powersupply voltage

Power per 0.77 kWthruster

Maximum thrust 8.9 mN

Specific impulse 7400 s

Propellant mass

Maximum in-space operationtime for onethruster

Longest groundtest

Comment

2_0 min.

1230 h

HV powersupply taileddue tocontaminationtiom gasesvented frombatteries. Ref.10

Electron Contactbombard- ioniza-ment lion

Mercury Cesium1 1

10 cm 8 cm

Ta wire Ta wire

2500 V 4500 V

1.4 kW 0.6 kW

28 mN 5.6 mN

4900 s 8050 s

31 min. 0 min.

1PS and Cesiumneutraliza- engine hadlion a HVdemonstra- short..lion. Ref. 3 Ref. 3

1 1

-7 cm -7 cm

Wire filament Wire filamentin beam

5000 V 5000 V

700

Contactionization

Cesium

1

-5 cm

Wirefilament.bariumcoated

4500 V

0.77 kW 0.77 kW -(I.4 kW

8.9 mN 8.9 mN

7400 s 7400 s

-19 min. -4 min.

Failed 3rd stageburn shortenedoperation,Obtained -20% offull thrust..Ref. 14

Stable operat ion of theIPS. S/C potential-1000 V at full thrust ..Ref. 14

-8.5 mN

5100 s


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Tableb. Experimentallightsof IonPropulsionystemsSpacecraft

Sponsor

Builder of IPS

Launch date

ATS-4 ATS-5

Orbitr km

IPS type

Propellant

No. of thrusters

Thruster anodediameter

Type ofneutralizer

Beam powersupply voltage

Power perthruster

Maximum thrust

SERT II ATS-6 SCATHAP78-2

USAF/NASA USAF/NASA NASA LeRC NASA GSFC USAF/NASAGSFC GSFC GSFCEOS EOS EOS

08.12,69

36,000

Contactionization

08.10.68

218x760

Contactionization

Cesium Cesium

2 2

5cmcm

Ta doped withYttrium

Ta doped withYttrium

3000 V 3000 V

0.02 kW 0.02 kW

0.089 mN 0.089 mN

Specific impulse 6700 s 6700 s

-0.05 k Eropellant massMax. in-space -10 h No operationoperation time withaHVfor one thruster beam.

Longest ground 2245 htest_s)

Comment

NASA LeRC,Westin[house02.03.70

1000

05.30.74

36,000

Hughes(ion source)

01.30.7943,000x27,000

S/C was in a lowaltitude parkingorbit due to aCentaur stagefailure. Firstsuccessfullorbital test of anion engine. NoEMI to S/Csubsystems.Ref. 18

Electron Electron Electronbombardment bombardment bombardment

Mercury Cesium Xenon2 2 1

15 cm 8 cm 3.6 cm

Hollow Cesiated Ta Ta doped withcathode Yttrium

3000 V 56O V

0.15 kW

1000 V to2OO0 V

0.03 to 0.045kW

4.5 mN 0.14 mN2500 s 350 s

3.6 k_ 0.3 k_92 h

-600 h

S/C had a 76RPM spin=rate.This produced a4g field whichcompromised thecesium feedsystem andprecludednormaloperation.Ref. 21

0.85 kW

28 mN4200 s

15 k_-3781 h

6742 h,5169 h

One ion engineoperated 3781 huntil theneutralizer tankwas depleted.The other enginehad a grid shortwhich limitedoperation to2011 h. Ref. 23

2614 h,471 c_/cles

One thrusteroperated for ~1hour and theother for 92hours. Furtherthrusting wasterminated due toa feed system"flooding"problem. NoEMI. Ref. 24

Operations wereperformedintermittantlyover a 247 dayperiod. Ref. 26

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Table2.MaiorIonPropulsionystememonstrationsProject SIT-5Name

Sponsor NASA LeRC

Builder of Hughesthruster

Builder ofPPU

IntegratorIPS

Project 1969 to 1972duration

Propellant Mercury

Thruater 5 cmdiameter

Hollow cathodeype ofneutralizer

Beam powersupplyvoltage

Power perthruster

Maximumthrust

Specificimpulse

Longestground test

SEPST

NASA JPL

JPL

Hughes/TRW

of JPL

1968 to 1972

Mercury

20 cm

Hollow cathode

2000 V

2.5 kW

88 mN

1600 V

0.072 kW

SEPS

NASA

Hughes

TRW

LeRC

1972 to 1980

Mercury

30 cm

Hol low cathode

I100 V

lAPS

NASA LeRC

Hughes

Hughes

Hughes

1974 to 1983

Mercury

8 cm

Hollow cathode

1200 V

XIPS-25

INTELSAT

Hughes

Hughes

Hughes

1985 to 1988

Xenon

25 cm

Hol low cathode

750 V

XIPS-25

Hughes

Hughes

Hughes

Hughes

Ongoing in1999. preflight

Xenon

25 cm

Hollow cathode

1400 V

3600 s

1300 h

2.1 mN

3000 s

2.6 kW

128 mN

3000 s

0.13 kW

5.1 mN

2500 s

1.3 kW

63 mN

2800 s

4.2 kW

165 mN

3800 s

9715 h 10,000 h 15,040 h,9489 h, 7112 h

4350 h, 3850cycles

Ongoing in1999

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Table3. OperationallightsofIonPropulsionystemsPAS-5 Galaxy ASTRA- Deep SATMEX PAS 6B

Spacecraft VlII-i 2A Space 1 5

Sponsor PanAmSat PanAmSat SES NASA Satmex PanAmSatLeRC/JPL

i

Builder of IPS Hu_hes Hu_hes Hu_hes Huhes Hu_hes Hu_hesLaunch date 08.27.97 12.08.97 08.29.98 10.24.98 12.05.98 12.21.98

Orbit, km 36.000 36,000 36,000 Orbits sun 36,000 36,000

IPS type Electron Electron Electron Electron Electron Electronbombardm't bombardm't bombardm't bombardm't bombardm't bombardm't

Xenon Xenon Xenon Xenon Xenon XenonPropellant

No. of thrusters 4 4 4 1 4 4

Thruster diameter 13cm 13cm 13cm 30cm 13cm 13cm

Type of Hollow Hollow Hollow Hollow Hollow Hollowneutralize r cathode cathode cathode cathode cathode cathode

Beam power 750 V 750 V 750 V 650 V to 750 V 750 Vsupply volta[[e 1100 V

0.44 kW 0.44 kW 0.44 kW 0.44 kW 0.44 kWower perthruster

0.50 kW to2.3 kW

Maximum thrust 18mN 18mN 18raN 92raN 18mN 18mN

Specific impulse 2590 s 2590 s 2590 s 1900 s to 2590 s 2590 s3100 s

Propellant mass >100 kg >100 kg >100 kg 82 kg >100 kg >100 kg

Maximum in- 1764 h as ofspace operation 04.27.99time for onethruster

Longest ground >8000 h 8193 htest

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Table 4. Comparison of the Propellant Throughput Capability of theSERT II and Deep Space I Ion Propulsion Systems

SERT II

Propellant type

_on_est ,ground test

Maximum thruster propellant throughput in-space

NSTAR/Dee 9 Snace 1

Propellant type

.onE.est ground test

Maximum thruster propellant throughput in-spaceas of 04.27.99

DS I propellant throughput capability for theprimary and extended mission

Propellant Throughput

Me_u_

-16 k_

Estimated to be -9 kg

Xenon

87.5 k_I 1.6 kg

82 kg

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Figure. Historyofelectron-bombardmentonthrusterevelopmentntheUnitedStates.(All projectsereNASAsponsorednlessotedtherwise.)YEAR1960 ->

1964 ->

1966 ->

1970 ->1972 ->

1973->

1976->

1980 ->1981->

1988->

1997 ->1998 ->

1999 ->

COMPONENT DEVELOPMENT _ADVANCES PROGRAMS

I Multi -aperture grids

VaporizerLong-life oxide maincathodeP lasma bridgeneutralizer anddischarge chamberhol low ca thode

I HV, propellantiso la tor {Hu_hes)

I Dished [zridsGrid erosion control

Control of spalledfl akes i n dischargechamberTest facility effects oncomponent wear

I Chan_e H_ -> XeI Rin_.-cusp chamber

Develop reliable Xehollow cathode viaSpace St at ion plasmacontactor program

10-cm lab thruster II-cm lab thruster20-cm lab thruster ]

SERT II EM thruster(15-cm)

50 -cm lab thrus ter150-cm lab thrust er

20-cm SEPST EMs_'stem5-cm EM thruster8-cm lab thruster30 -cm lab thrus ter30 cm EMDevelopment Cont ractat Hu_hesSEPS developmentpro[zram8-cm EM thruster

lAPS developmentprogram (8-cm, H_)

30-cm thruster (Xe)25-cm thruster (Xel([NTELSAT/Hu_hes)13-cm lab thruster (Xe)(Hu_hes)

] SERTI (10-cm)

SERT II thruster& PPU I 15-cm SERT 11_,ound-tested 6742 h I f light svstem

15,_ h test-8 cm

I 0,000 h test - 30 emEM5070 h test-30 cm EM

XIPS-25 for comsatorbit insertion andNSSK (Hu_hes)Initiate development ofsubkilowatt and 5 kWIPS for Earth-orbitaland deep space S/C

9489 h test of the 8-cm. IEM mercury thruster I

>8000 h test ofXIPS-13 (Hu_hes)8193 h test of theNSTAR thrusterExtended testing of theXIPS-25 (Hughes)

XIPS- 13 for comsatNSSK (Hu_hcs)

NSTAR 30-cm lorDeep Space 1

Ext ended ground-testing of the NSTARf light spa re thrus ter ,PPU, and DCIU

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Figure 2. Ion propulsion technology roadmap for Earth-orbital and planetary applications

(Calendar Year)

1995

Hughes PAS-5,500 W XIPS(First of many S/C)

2000 2005 2010t t 0Hughes 701S/C4.2 kWXIPS

INSTAR on. I[eep Space I,2.5 kW |PS

Flight Applications

Space Technology 4,3 NSTARsubsystem IPS

2 X CostReductionin IPS

1Taurus Planetary mission, 1Europa Orbiter, Outer Planet Explorer_li_i_High Delta-V Orbit Transfers, _"DoD Missions /

t2.5 kW IPS

Core Technology

5 kW 1PS High Power IPSI Low Power IPS

JCathode Low FlowTechnologies Neutralizer

1Titanium Carbon- Sub-Kilowatt Sub 100 _' 10 kW- 30 k_'Ion Optics Carbon Thruster Thruster Systems Fon Optics (Efficiency) (low mass)

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REPORT DOCUMENTATION PAGE Fo_ApprovedOMB No. 0704-0188Public reporting burden for this collection of information is estimated to average 1 hour per response, incl uding the lime for reviewing instructions, searching existing data sources.gathering and maint aining the data needed, and completing and rev iewing the collection of information. Send comments regarding this burden estimate or any other aspect of thiscollection of information, including suggestions for reducing this burden, Io Washinglon Headquarters Services, Directorate for information Operations and Reports, 1215 Jeff ersonDavis Highway, Suite 1204. Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 205031. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

October 1999 Technical Memorandum4. TITLE AND SUBTITLE

A Synopsis of lon Propulsion Development Projects in the United States:SERT I to Deep Space I

6. AUTHOR(S)James S. Sovey, Vincent K. Rawlin, and Michael J. Patterson

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)National Aeronautics and Space AdministrationJohn H. Glenn Research Center at Lewis FieldCleveland, Ohio 44135-3191

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space AdministrationWashington, DC 20546-0001

5. FUNDING NUMBERS

WU-632-1B-I B-00

8. PERFORMING ORGANIZATIONREPORT NUMBER

E-11916

10. SPONSORING/MONITORINGAGENCY REPORTNUMBER

NASA TM-- 1999-209439AIAA 99-2270

11. SUPPLEMENTARY NOTESPrepared for the 35th Joint Propulsion Conference and Exhibit cosponsored by the AIAA, ASME, SAE. and ASEELos Angeles, California, June 20-24, 1999. Responsible person, James S. Sovey, organization code 5430,(216) 977-7454.

12a. DISTRIBUTION/AVAILABILITY STATEMENTUnclassified - UnlimitedSubject Category: 20 Distribution: NonstandardThis publication is available from the NASA Center for AeroSpace Information. (301) 621-0390.

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)The historical background and characteristics of the experimental flights of ion propulsion systems and the major ground-based technology demonstrations were reviewed. The results of the first successful ion engine flight in 1964, SERT I whichdemonstrated ion beam neutralization, are discussed along with the extended operation of SERT II starting in 1970. Theseresults together with the technology employed on the early cesium engine flights, the ATS series, and the ground-testdemonstrations, have provided the evolutionary path for the development of xenon ion thruster component technologies,control systems, and power circuit implementations. In the 1997-1999 period, the communication satellite flights using ionengine systems and the Deep Space 1 flight confirmed that these auxiliary and primary propulsion systems have advancedto a high-level of flight-readiness.

14. SUBJECT TERMSPropulsion; Electric propulsion; Spacecraft propulsion; Plasma applications

17. SECURITY CLASSIFICATIONOF REPORT

UnclassifiedNSN 7540-01-280-5500

18. SECURITY CLASSIFICATIONOF THIS PAGEUnclassified

19. SECURITY CLASSIFICATIONOFABSTRACTUnclassified

15. NUMBER OF PAGES25

16. PRICE CODEA03

20. LIMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)Prescribed by ANSi Std. Z39-1 8298-102


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A Synopsis of Ion Propulsion Development Projects in the US SERT 1 to Deep Space I