NASA/TMm1999-209447 AIAA-99-0457

Performance Evaluation of a 50 kW

Hall Thruster

David T. Jacobson

Dynacs Engineering Company, Inc., Brook Park, Ohio

Robert S. Jankovsky

Glenn ResearchCenter, Cleveland, Ohio

December 1999

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NASA / TMm1999-209447 AIAA-99-0457

Performance Evaluation of

Hall Thruster

a 50 kW

David T. Jacobson

Dynacs Engineering Company, Inc., Brook Park, Ohio

Robert S. Jankovsky

Glenn Research Center, Cleveland, Ohio

Prepared for the

37th Aerospace Sciences Meeting and Exhibit

sponsored by the American Institute of Aeronautics and Astronautics

Reno Nevada, January 11-19, 1999

National Aeronautics and

Space Administration

Glenn Research Center

December 1999
Acknowledgments

The authors would like to thank Mr. Thomas Haag and Mr. Luis Pinero of NASA GRC for their technical

assistance. A sincere thanks is also extended to Kevin Blake, Wanye Condo, Eli Green, George Jacynycz,

Randy Mele, and John Miller of NASA's Test Installation Division for their contribution.

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

Price Code: A03

Available from

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

Price Code: A03
PERFORMANCEEVALUATIONOFA50KWHALLTHRUSTER

DavidT.Jacobson

DynacsEngineeringCo.,Inc.

BrookPark,Ohio44142

RobertS.Jankovsky

NationalAeronauticsandSpaceAdministration

GlennResearchCenter

Cleveland,Ohio44135

ABSTRACT

An experimental investigation was conducted on a laboratory model Hall thruster designed to operate at

power levels up to 50 kW. During this investigation the engine's performance was characterized over a

range of discharge currents from 10 to 36 A and a range qf discharc;e voltages fronl 200 to 800 V.

Operating on the Russian cathode a maximum thrust of 966 mN was nleasured at 35.6 A and 713.0 V.

This corresponded to a spec!fic impulse of 3325 s and an efficiency of 62%. The maximum power the

engine was operated at was 25kW. Additional testing was conducted using a NASA cathode designed for

higher current operation. During this testing, thrust over 1 N was measured at 40.2 A and 548.9 V.

Several issues related to operation qf Hall thrusters at these high powers were encountered.

Introduction

Beginning in 1990, NASA initiated an effort

to evaluate Hall thrusters due to their

unique combination of relatively high

specific impulse and favorable thrust

density. These efforts primarily focused on

the Russian Stationary Plasma Thruster

(SPT) and Thruster with Anode Layer

(TAL). These engines represent technology

developed to a relative state of maturity

over several decades in the former Soviet

Union. As an example, over 70 SPT's have

been used operationally on Soviet and

Russian spacecraft since 1971 _. The anode

layer thruster, which potentially offers

higher performance than SPT type thrusters,

was never flown on a Soviet or Russian

spacecraft, but has now been demonstrated

on-board a Western experimental spacecraft

with its first successful firing on October 23,

1998-'.

space power beaming satellite _. In order to

validate the engine technology needed for

such missions at the high total powers being

considered, NASA's Advanced Space

Transportation Program has undertaken a

demonstration of a high power Hall thruster

capable of operation at power levels

approaching 50 kW.

The objective of this demonstration was to

assess the suitability of this approach and to

gain insight into the relevant issues with

operating high power Hall thrusters of this

type. The performance of one high power

engine was experimentally evaluated during

this investigation. These results are

included in this paper for a wide range of

operating conditions. The problems

encountered during these tests are also

discussed.

In recent years with advances in space

power systems ) and the consideration of

new orbital trajectories 4, government

mission designers have begun reevaluating

the potential of leveraging these

technologies to reduce launch mass for

various potential missions. These missions

include a manned mission to Mars _ and a

Apparatus and Procedure

The Hall thruster used in this investigation,

referred to as the TM-50, was a laboratory

model anode layer thruster obtained from

the Central Research Institute of Machine

Building (TsNIIMASH). This engine was

developed several years ago to evaluate the

NASA / TM--1999-209447 1
concept of applying Hall thruster

technologyat the tensof kilowatt power

level, but was never tested to its full

designed capabilities due to facility

constraintsL The anticipatedmaximum

input powerfor thisenginewas50kW. A

photographof the hardwareis shownin

Figure1.

The device had an averagedischarge

chamberdiameterof 200mmand overall

dimensionsof 310mm x 310ram x 160 mm.

The thruster mass was approximately 30.5

kg. The thruster employed a conductive

discharge chamber as do all anode layer

thrusters. In this case, however, the

discharge chamber was segmented into two

stages allowing for operation in various

electrical configurations. The first

configuration, termed single stage was

achieved by floating the second stage anode.

An alternative single stage configuration

may be achieved by electrically shorting the

two anodes. Referring to Figure 2a, these

different electrical configurations are

achieved by placing the switch in position A

or B respectively. Another configuration,

shown in Figure 2b that was briefly

investigated was two stage operation. In

this configuration, each segment is powered

with a separate power supply as discussed

in more detail later.

The cathode is mounted in the center of

thruster along the thruster axis. Initial

testing was completed using a Russian

hollow cathode that was delivered with the

thruster. This cathode was designed to

operate at current levels up to 20 A with a

xenon propellant flow rate from 1-3 mg/s.

Subsequent testing at higher power levels

was conducted with a NASA hollow

cathode that has higher emission current

capabilities. This cathode is based on the

Plasma Contactor design originally

developed for the International Space

Station. _

The thrust stand used in these tests was

originally designed for a 0.5 MW

magnetoplasmadynamic (MPD) thruster. _ A

photograph of the thrust stand is provided

as Figure 3. The thrust stand held the

thruster with its horizontal axis

approximately coincident with the

centerline of the vacuum facility. The stand

was an inverted pendulum type design

which allowed up to 25 mm of travel along

the sensitive axis. An elastic restoring force

was provided by the deformation of a load

spring as well as the support flexures. This

gave the system a natural frequency of

approximately 0.25 Hz. Thrust induced

deflections were measured using a linear

variable differential transformer (LVDT)

and recorded on a strip chart recorder.

Transient motion of the thrust stand was

critically damped with an electronic

feedback circuit acting through an

electromagnetic actuator. The damper only

applied force when motion was detected,

and did not affect steady state thrust

measurements.

Xenon gas for the cathode and main

discharge was transferred onto the thrust

stand using two separate propellant

flexures. These consisted of 3ram diameter

stainless steel tubes with one end anchored

to the stationary facility, and the other end

attached to the moving structure of the

thrust stand. Electrical current for the

cathode and anode was transferred through

the stand to the thruster using water cooled

electrical flexures. The conductive spans

were isolated at each end with plastic

tubing. Electrical currents for the cathode

keeper, heater, and electromagnets were

transferred through the thrust stand, to the

thruster with multi-conductor ribbon cables.

The thrust stand was calibrated in place and

under vacuum. A mono-filament nylon line

was attached to the rear of the thruster and

aligned with the thrust vector. The line

passed over a pulley and was drawn by

weights which could be cycled on and off.

The thrust stand response to the calibration

weights was recorded on the strip chart.

This calibration could then be used to

interpret the displacement measured during

thruster operation. Cycling the calibration

weights revealed hysterisis on the order of

1% of the full scale value. Zero drift was

also observed over periods in which the

thruster was fired. The amount of drift was

NASA / TM--1999-209447 2
dependentonthepowerlevelandduration

of the test. Sincethe drift ratewasfairly

consistent,thrust datawere correctedfor

thiserror.

CommerciallyavailableDCpowersupplies

wereusedto operatetheengine.Separate

supplieswereusedto powertheinnerand

outermagnets,cathodeheater,keeperand

ignitor. Two60A,500Vdischargesupplies

configuredin serieswere usedfor single

stageoperation.A lowvoltage,highcurrent

keepersupplyprotectedbyablockingdiode

wasusedin parallelwith a high voltage,

low current ignitor supply for cathode

ignition and operation.An output filter

consistingof a3.8fl ballastresistorin series

with the anode and 100 _tF capacitor

betweencathodeandanodewasused.Back

to back75volt Zenerdiodeswereusedto

clampthefloatingpotentialof thecathode

withrespectto ground.

All performancedatareportedhereinwere

takenwith the thrusteroperatingin the

singlestageconfiguration. A schematic

diagramof theelectricalconfigurationused

for the investigationis providedasFigure

2a. Thethrusterwasoperatedin the two

stagearrangementto verifytheintegrityof

the power system. Electrically this

configurationdiffered from single stage

operationby placing each of the two

separate60 A, 500 V dischargepower

supplieson a different segmentof the

anode. This configuration is shown

schematicallyinFigure2b.

Datawere collectedusinga commercially

available data acquisition system.

Dischargecurrent and voltage, magnet

current and voltage, floating potential,

keepercurrent,anodeand cathodeflow

rates, and engine temperatureswere

monitored.Dataweresampledat a rateof

1Hz. Simple internal mathematical

functionsallowedpreliminarycalculationof

power, thrust, specific impulse, and

efficiency.

A laboratoryfeed systemwas used to

providepropellanto theanodeandcathode

at the desired mass flow rates via

commerciallyavailableflow controllers.

Theaccuracyofthesedevicesquotedbythe

manufacturerwas_+1.0%full scale.Flow

calibrationswere performedbefore and

after testing using a constant volume

method.Commerciallyavailableresearch

gradexenonwith a purity of 99.999%was

used. The feedsystemwashelium leak

checked prior to operation of the

propulsionsystem.

All testing was conductedin the main

volumeofVacuumFacility5atNASAGRC.

A detaileddescriptionof thefacilitycanbe

foundin ref.9. Thischamberis 4.6m in

diameterand19m in length.It isequipped

with20diffusionpumpswhichare0.8m in

diameter. This facility pumps xenonat

approximately100,0001/s with only the

diffusionpumpsoperating.Thefacility is

alsoequippedwith a gaseousheliumcryo

systemwhichpumpsxenonatanadditional

1.5million1/s.A photographof thefacility

is providedasFigure4. The facilitywas

initially operatedonly on diffusionpumps

to facilitaterapid facility turnaround if

problemsweredetectedwith thehardware

or testsetup. Datawerecollectedfor low

propellantflowoperatingpoints.As engine

flow rateswereincreased,thecryosystem

was started. The highesttank pressure

experiencedduringtestingwas5.3x10_torr

measuredusingionizationgaugeslocated

on the facilitywall, calibratedon air and

uncorrectedforxenon.

Followingthefacilitypumpdownsequence,

the enginewas permitted to remain at

vacuumfor 24hoursprior to operationto

allow for outgassingof the hydroscopic

componentsof theengine.Thisprocedure

was recommendedby the manufacturer

TsNIIMASH. The current limits of the

dischargepowersuppliesweresetto 25A

prior to operation.TheRussiancold start

cathodeprovided with the engine was

ignitedby applyinghigh voltagebetween

thecathodeand keeper. Thexenonmass

flowratewassetto 3mg/s forignitionand

1.5mg/s for steadystateoperation. The

inner and outer magnet supplies were

switchedon with currentvaluesof 1.0A

and 0.5 A respectively.With the cathode

NASA/TM--1999-209447 3
ignited,magnetsenergized,and an anode

flow rateof 10mg/s, the main discharge

wasstartedby rampingup the discharge

voltageuntil breakdownoccurred. The

innermagnetcurrenttypicallywasreduced

fromthesuggestedvalueof 1.0A to0.5A to

facilitatebreakdownat a lower discharge

voltagevalue. With the reducedinner

magnet current, breakdownconsistently

occurredatadischargevoltagelessthan100

V. Thedischargevoltagewasincreasedto

300V overa periodof 0.5hoursin 50 V

increments to allow for hardware

outgassing. The power level was increased

to 10 kW by increasing the discharge current

and voltage values in a stepwise fashion.

This procedure was followed to allow the

thruster to reach thermal equilibrium. Prior

to making performance measurements, the

engine was shut down and a thrust

calibration was performed as described

previously.

During performance evaluations the engine

was operated by setting the mass flow rate

to the desired value and varying the

discharge voltage. When oscillations in the

discharge current created thruster

instabilities, the inner and outer magnets

were optimized by attempting to minimize

the magnitude of the discharge current.

Results and Discussion

During this investigation, performance data

were taken with the engine operating in the

single stage electrical configuration at power

levels ranging from 2.0 to 25.4 kW. At the

highest power level of 25.4 kW, the engine

produced 966 mN of thrust with specific

impulse of 3325 s and efficiency 62%. At an

alternate operating point, thrust over 1 N

was achieved with specific impulse of 2936 s

and 66% efficiency. Several issues regarding

testing this high power engine were

encountered.

Performance

The majority of data discussed herein were

taken with the engine operating on the

Russian cathode provided. Because the

device was not optimized for operation with

this Hall thruster, specific impulse and

efficiency values were calculated without

the mass flow of the cathode. Also, the

efficiencies reported were calculated using

the discharge power only and did not

include the power dissipated in the magnets

(typically less than 10 W).

The relationship between thrust and power

is shown for eight mass flow rates in Figure

5. The measured thrust values reported

ranged from 147 mN to 966 mN. Thrust of

1011 mN was achieved at the power level of

22 kW with specific impulse of 2936 s and 66

% efficiency (this point was taken using the

NASA cathode). While a comprehensive

tabulation of those results taken with the

NASA cathode is not presented here, a table

of the performance test results taken with

the Russian cathode is provided in Table 1.

Discharge current is plotted as a function of

anode mass flow rate in Figure 6. Discharge

current is typically directly proportional to

anode mass flow rate. In the case of the TM50 operating at high mass flow rates, the

relationship deviates from linear. From this

plot, it can be seen that the discharge current

is significantly higher for a given mass flow

rate; for example, at the highest mass flow

rate of 29.6 rag/s, a discharge current value

of 35.6 A was measured. Also, from the plot

it can be seen that there is a vertical spread

of the discharge current values for a given

mass flow rate. This behavior was

attributed to operation with a nonoptimized magnetic field. As noted

previously, magnets were adjusted such that

the discharge current was minimized. This

procedure was not implemented at each

data point because identification of the

optimum performance points was not the

objective of this test sequence given the time

constraints of this investigation.

Specific impulse and efficiency versus

thruster power are plotted for the eight

mass flow rates tested in Figure 7 and 8

respectively. Specific impulse increased

with voltage in nearly all cases. Cases do

exist were a increase in voltage did not yield

an increase in specific impulse, but in those

instances, there was a correspondingly

sharp drop in thruster efficiency. This

NASA / TM--1999-209447 4
behaviorwas attributedto non-optimized

magnetswhichcauseda modalchangein

thrusteroperation.Thetrendof increasing

specificimpulsewith voltagedid comeat

theexpenseofefficiencyasdiscussedbelow.

Efficiency versus dischargevoltage is

plottedfor eightanodemassflow ratesin

Figure 9. Efficiency increasedwith

dischargevoltageup to approximately500

V at which point the efficiencybeganto

decreaseas the voltage was increased

further.This behaviorexistsfor all mass

flow ratesat whicha dischargevoltageof

500V wasexceeded.Asa result,operation

of largeHall thrustersof thistypeover500

V does not seemadvantageousfrom a

performanceperspective.However,it may

be possibleto obtain high efficiencyat

dischargevoltages>500Vby operatingin a

twostageconfiguration.

When operating in the two stage

configuration,thedischargeisinitiatedwith

theacceleratingsupplybetweenthecathode

andanode2,whichisthesegmentclosestto

the exit plane. The dischargesupply

betweenanode1 and anode2 is then

energized. Application of accelerating

voltagesbetween500V and2000V allows

for operationin ahighspecificimpulseand

efficiencymodewith relativelylow thrust.

Anefficiencyofover70%wascalculatedfor

datatakeninRussiawhileoperatingwithan

anode mass flow rate of 10.16rag/s,

dischargevoltage of 330 V, and an

acceleratingvoltage900V_'.Relativelyhigh

thrustwith correspondinglylowerspecific

impulseandefficiencycanbeachievedby

operating in the single stage mode.

Utilizing the favorablecharacteristicsof

eachmodeof operation,atwostagethruster

couldbeusedto performmultiplemission

requirements.Only a shortdurationtest

sequencewasconductedin the two stage

configurationat NASA GRC to verify

proper operation of the engine and

laboratorypowersystem.

Testin_g

In general,the thrusteroperatedwithout

incident at power levels below 10 kW.

Currentoscillationswereeasilyminimized

by adjustingthe magneticfield. Engine

breakdownoccurredreadilyas described

above. However,at elevatedpowerlevels

(>10kW)anomalousarcingbetweenanode

and thrusterbody periodicallyoccurred.

Specifically,arcingoccurredbetweenanode

2 and the graphiteguard rings which

protectthemagneticsystem.Therewerea

very small numberof arc occurrencesat

powerlevelsbelow10kWbutthebehavior

wasmoreprevalentatvoltagesabove400V

andanodemassflow ratesabove25mg/s.

As a result no performancedata were

obtainedfor powersabove25kW.

Severalfactorswereidentifiedas possible

influenceson the arc occurrences: the

temperatureof theengine,facilitypressure,

internal enginematerials,magneticfield

effects,cathodeflowandposition,capacitor

size,temporalresponseof the discharge

power supplies, or an anomalous

connectionin theelectricalscheme.Several

of these items have been investigated

withoutconclusivevidenceof a significant

impacton the engine'soperation. It has

beendemonstratedthat sustainedversus

unstained arc occurrences have a

dependenceon the valueof thedischarge

currentlimit; forexample,sustainedarcsdo

notoccurreadilywith thedischargecurrent

limit set below20 A. A comprehensive

evaluationis currently underwayin an

attempttomorecompletelyunderstandthis

behavior.

Severalissuespertinento thetestapparatus

becameapparentwhile testing at high

powerlevels.Ahightemperaturelastomer

tubingwasusedto makethe final anode

connectionsbetweenthethrusterandthrust

stand. This tubing wasused to provide

electricisolationof theanode.Whiletesting

at highpowers,oneof thethreetubeswas

damagedasaresultof anover-temperature

condition.Thisin turncauseda propellant

leakin theanodeline. Forsubsequent tests,

the distancebetweenthe thrusterand the

elastomerto metaltubing connectionwas

increasedto reducethetemperatureat the

isolatorinterface.

NASA/ TM--1999-209447 5
Additionalelectricalisolationissueswere

alsomanifestduringhighpowertesting.For

thistestseries,electricalconnectionsinside

the vacuum chamberwere made using

mechanicalfasteners. Siliconetapewas

used at the connectionfor electrical

isolation. Thismethodhasbeenusedin

previous Hall thruster investigations

without consequence. Problemswere

experiencedin maintainingtheintegrityof

theelectricalisolationon the high current

connections.Therewereseveralinstancesof

electricalarcsto facility groundat these

locations. Additional plastic insulation,

fiberglassleeving,andsiliconetubingwere

addedin subsequentteststo mitigatethese

problems.

Concluding Remarks

An investigation was conducted to assess

the performance of a high power Hall

thruster. Performance measurements were

made at power levels ranging from 2.0 to

25.4 kW. At the highest power level, the

engine produced 966 mN of thrust with

specific impulse 3325 s and efficiency of

62%. Higher power levels were not

achieved due to arc occurrences between the

anode of the second stage and the thruster

guard rings. An investigation is currently

being conducted to determine the cause of

the arcing phenomena. Additional

performance measurements will be taken at

power levels in excess of 25 kW for the

single stage configuration after the arcing

issues have been resolved.

All data reported were taken with the

thruster operating in the single stage

configuration. An extensive investigation of

the thruster's two-stage performance will be

conducted in the future.

1.

.

.

.

.

.

.

,

9.

References

Bober, A.S. and Maslennikov, "SPT in

Russia-NewAchievements,"

Proceedings of the 24 '_ International

Electric Propulsion Conference, Sept.

1995, pp. 54-60.

Naval Research Laboratory, "Naval

Research Laboratory Goes Operational

With United States First Hall Thruster

Electric Propulsion System," Press

Release, October 29, 1998.

Bennet, G. "Aerospace Power", The Year

in Review, Aerospace America,pg. 58,

Dec. 1998.

Reference Mission Version 3.0

Addendum to the Human Exploration

of Mars: The Reference Mission of the

NASA Mars Exploration Study Team.

Ex13-98-036, June, 1998, NASA Johnson

Space Center, Houston, TX 77058.

Oleson, S.R., White Paper: "Electric

Propulsion Technology Projections for

Space Solar Power" NASA Lewis

Research Center, Nov. 13, 1998.

TsNIIMASH-Export, "TM-50 Thruster

with Anode Layer Operations Manual,"

Internal TsNIIMASH document, 1998.

Patterson, M.J., Hamley, J.A., SarverVerhey, T.R., Soulas, G.C., Parkes, J.E.,

Ohlinger, W.L., Schaffner, M.S., and

Nelson, A., "Plasma Contactor

Technology for Space Sation Freedom,"

AIAA-93-2228, June 1993.

Haag, T.W., "Thrust stand for highpower electric propulsion devices," Rev.

Sci. Instrum. Vol. 62, No. 5, pp. 1186-

1191, May 1991.

Sovey, J.S., Vetrone, R.H., Grisnik, S.P.,

Myers, R.M., Parkes, J.E. "Test Facilities

for High-Power Electric Propulsion,"

Journal of Propulsion and Power, Vol.

10 No.l, Jan. 1994.

NASA/TM---1999-209447 6
Table1:TM-50Hallthrusterperformancedata.

Discharge

Voltage

V

Discharge

Current

A

Power

W

Anode

Flow

mg/s

Cathode

Flow

mg/s

Thrust

mN

Efficiency Specific

Impulse

197.8 10.3 2037 10.3 1.5 147 0.52 1454

200.5 10.2 2045 10.3 1.5 147 0.51 1456

250.6 10.4 2606 10.3 1.5 169 0.53 1673

299.6 10.2 3056 10.3 1.5 176 0.49 1741

351.9 10.0 3519 10.3 1.5 189 0.49 1866

401.4 9.9 3974 10.3 1.5 202 0.50 1993

9.9 10.3 1.5

1.5

1.5

15.0 14.9

213

255

276

4444

4473

0.49

0.49

0.49

448.9

298.2

2102

1748

348.7 15.0 5231 14.9 1890

401.3 15.0 6020 14.9 1.5 305 0.52 2094

6728 14.9 1.5 324 2222

1.5

15.0

17.2

448.5

149.8

0.53

2577 16.3 191 0.44 1198

199.3 17.6 3508 16.3 1.5 239 0.50 1494

249.9 17.0 4248 16.3 1.5 281 0.57 1759

300.0 16.8 5040 16.3 1.5 313 0.60 1958

350.0 16.7 5845 16.3 1.5 336 0.59 2103

401.2 16.6 6660 16.3 1.5 358 0.59 2241

451.3 16.7 7537 16.3 1.5 376 0.58 2354

501.9 16.7 8382 16.3 1.5 395 0.57 2475

501.5 16.6 8325 16.3 1.5 390 0.56 2445

552.0 16.6 9163 16.3 1.5 407 0.56 2551

549.9 16.5 9073 16.3 1.5 411 0.57 2574

602.0 16.6 9993 16.3 1.5 429 0.57 2686

650.3 16.6 10795 16.3 1.5 441 0.55 2763

704.5 16.6 11695 16.3 1.5 456 0.55 2854

758.7 16.8 12746 16.3 1.5 470 0.53 2942

13472 16.3 1.5 484

18.4

16.7

20.6

0.53

5539 1.5 336 0.55

806.7

268.9

3030

1862

309.6 20.5 6347 18.4 1.5 361 0.56 2000

349.3 20.4

149.7

388 0.57

255 0.49

18.4

20.2

7126

22.2 3323

1.5

1.5

2150

1289

200.8 22.0 4418 20.2 1.5 314 0.55 1585

251.0 21.9 5497 20.2 1.5 360 0.58 1819

302.9 22.4 6785 20.2 1.5 409 0.61 2067

350.9 22.8 8001 20.2 1.5 443 0.61 2238

400.6 22.7 9094 20.2 1.5 474 0.61 2392

449.6 22.7 10206 20.2 1.5 500 0.61 2524

501.1 22.6 11325 20.2 1.5 528 0.61 2669

502.6 21.5 10806 20.2 1.5 516 0.61 2605

NASA/ TM--1999-209447 7
Table1:continued

Discharge

Voltage

V

Discharge

Current

A

Power

W

Anode

Flow

mg/s

Cathode

Flow

mg/s

Thrust

mN

Efficiency Specific

Impulse

s

276.6 24.1 6666 21.5 1.5 403 0.57 1912

300.5 23.9 7182 21.5 1.5 424 0.58 2014

353.0 23.7 8366 21.5 1.5 462 0.59 2193

402.4 23.5 9456 21.5 1.5 495 0.60 2349

452.0 23.4 10577 21.5 1.5 528 0.61 2505

450.9 23.2 10461 21.5 1.5 528 0.62 2505

500.8 23.1 11568 21.5 1.5 557 0.62 2641

549.2 23.2 12741 21.5 1.5 583 0.62 2769

598.0 23.1 13814 21.5 1.5 603 0.61 2862

650.5 23.2 15092 21.5 1.5 627 0.61 2975

700.0 23.0 16100 21.5 1.5 623 0.56 2955

745.5 23.0 17147 21.5 1.5 662 0.59 3139

351.6 27.5 9669 24.7 1.5 550 0.63 2273

402.9 27.0 10878 24.7 1.5 594 0.66 2455

455.6 26.8 12210 24.7 1.5 636 0.67 2628

500.0 26.6 13300 24.7 1.5 661 0.66 2729

548.2 26.6 14582 24.7 1.5 695 0.67 2871

600.0 26.8 16080 24.7 1.5 726 0.66 2998

650.1 26.9 17488 24.7 1.5 753 0.66 3109

697.6 26.9 18765 24.7 1.5 775 0.65 3200

748.2 27.3 20426 24.7 1.5 802 0.64 3311

449.2 27.4 12308 24.8 1.5 624 0.64 2567

501.9 27.4 13752 24.8 1.5 655 0.63 2695

350.0 28.3 9905 24.7 1.5 557 0.63 2295

301.4 29.8 8982 24.7 1.5 509 0.58 2099

357.8 34.8 12451 29.6 1.5 668 0.60 2299

399.2 33.9 13533 29.6 1.5 711 0.63 2446

453.7 33.7 15290 29.6 1.5 764 0.64 2630

509.5 33.4 17017 29.6 1.5 817 0.66 2814

551.3 33.3 18358 29.6 1.5 855 0.67 2942

600.3 33.3 19990 29.6 1.5 891 0.67 3066

649.5 33.3 21628 29.6 1.5 929 0.67 3199

713.0 35.6 25383 29.6 1.5 966 0.62 3325

NASA/TM--1999-209447 8
Figure 1:TM-50 Hall thruster with centrally mounted Russian cathode.

A

B

÷

Disch_ge

Supply

B m m i|m

Ignitor Keeper Heater

Supply Supply Supply !

I

4- 4"'q E ÷_

+1

I InnerMagnet Suppy

-- I

...I- I

Iouter Magnet Suppy

-- I

Figure 2a: Electrical schematic for the single stage configuration.

NASA / TM--1999-209447 9
8,8_

arge

ly

lOaF

Anode 1

Accelerating t

Supply

1001xF

Anode 2

m

I Ignitor Supply

÷1

Keeper Suply I Heater Supply

+_ + L_.._

IlnnerMagnet

su_pp,y ÷' II 1

+ I

OuterMagnetl <J

Supply I

-i _,

Figure 2b: Electrical schematic for two stage configuration.

Figure 3:TM-50 mounted on thrust stand in Vacuum Facility 5 at NASA GRC.

NASA/TM---1999-209447 10
Figure 4: Vacuum Facility 5 at NASA GRC.

1200

Z

"o

"O

e-.

r1000

8OO

600

400

200

A

A X

x

A x

Z_x x

_ m

X m"

o10.3 mg/s

'_'*-- 0o 14.9 mg/s

_XO.. []

._e _ .16.3 +18.4 mg/s mg/s

x

o_o a x20.2 mg/s -21.5 mg/s

o_o°° x24.7 mg/s

429.6 mg/s

I i I I I

0 5000 10000 15000 20000 25000 30000

Discharge Power (W)

Figure 5: Thrust produced by the TM-50 versus discharge power for eight different anode flow

rates.

NASA/TM--1999-209447 11
40.0

<

t_

35.0

30.0

25.0

20.0

15.0

10.0

5.0

Figure 6:

0

O

O

! ! ! i ! |

5.0 10.0 15.0 20.0 25.0 30.0

Anode Mass Flow Rate (mg/s)

Discharge current versus anode mass flow rate.

35.0

Figure 7:

3500

E

U

O

U3

3000

2500

2OOO

1500

1000

X

X t.

O "_ A

O --×

A

O --

°x_;'x x'< a

_X X A

m

A

o _ x o10.3 mg/s

o o +" a 14.9 mg/s

o _ o16.3 mg/s

o <u + 18.4 mg/s

O

x x 20.2 mg/s

o o -21.5mg/s

x 24.7 mg/s

I

o a29.6 mg/s

i i l I #

0 5000 10000 15000 20000 25000 30000

Thruster Power (W)

Specific Impulse versus thruster power for eight different anode flow rates.

NASA/TM--1999-209447 12
0.70

0.65

0.60

0.55

0.50

0.45

0.40

O

O

x

x

x

Yx x

m m

m

x _ x_x_x _ -

4>00 m

X -- X

o _°oo

" 0 0

X ++ 0 0

0

O0

[]tl

x

x_

x

x

o 10.3 mg/s

[] 14.9 mg/s

4>16.3 mg/s

+18.4 mg/s

x20.2 mg/s

-21.5 mg/s

x24.7 mg/s

a29.6 mg/s

i I i ! i

0 5000 10000 15000 20000 25000 30000

Thruster Power (W)

Figure 8: Thruster efficiency versus thruster power for eight different anode flow rates.

0.70

0.65

0.60

0.55

>,

.- 0.50

_2

r,1

0.45

0.40

0.35

0.30

x

O

o

x _

x _ _ x

x a _x

m -- mm

x_x_x -

x

o. + °_o o

-I-+ o

0

[] M

0 0 o o

x

x

A

o o

o 10.3 mg/s

[] 14.9 mg/s

o 16.3 mg/s

+ 18.4 mg/s

x 20.2 mg/s

-21.5 mg/s

×24.7 mg/s

A 29.6 mg/s

! i i l

0 200 400 600 800 1000

Discharge Voltage (V)

Figure 9: Thruster efficiency versus discharge voltage for eight different anode flow rates.

NASA/TM--1999-209447 13
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

Public reporting burden for this collection of informalion is estimated 1o average 1 hour per response, including the time for rewewmg _nstructions, searching existing data sources,

gathenng and maintaining the data needea, and completing and reviewing the collection of information Send comments regarding this burden estimate or any other aspect of this

collection of information including suggestions for reducJng this burden, to Washington HeadQuarters Services, Directorate for information Operations and Reports, 1215 Jefferson

Davis Haghway. Suite 1204, Arlington. VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE

December 1999

4. TITLE AND SUBTITLE

Performance Evaluation of a 50 kW Hall Thruster

6. AUTHOR(S)

David T. Jacobson and Robert S. Jankovsky

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135-3191

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

National Aeronautics and Space Administration

Washington, DC 20546-0001

3. REPORT TYPE AND DATES COVERED

Technical Memorandum

5. FUNDING NUMBERS

WU-632-1B-IB-00

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-11939

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM--1999-209447

AIAA-99-0457

11. SUPPLEMENTARY NOTES

Prepared for the 37th Aerospace Sciences Meeting and Exhibit sponsored by the American Institute of Aeronautics and

Astronautics, Reno Nevada, January l l-I 9, 1999. David T. Jacobson, Dynacs Engineering Company, Inc.. Brook Park,

Ohio, 44142, and Robert S. Jankovsky, NASA Glenn Research Center. Responsible person. Robert S. Jaknovsky,

organization code 5430. (216) 977-7515.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified- Unlimited

Subject Category: 20 Distribution: Nonstandard

This publication is available from the NASA Center for AeroSpace Information. (301) 621-0390

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

An experimental investigation was conducted on a laboratory model Hall thruster designed to operate at power levels up

to 50 kW. During this investigation the engine's performance was characterized over a range of discharge currents from

10 to 36 A and a range of discharge voltages from 200 to 800 V. Operating on the Russian cathode a maximum thrust of

966 naN was measured at 35.6A and 713.0 V. This corresponded to a specific impulse of 3325 s and an efficiency of

62q. The maximum power the engine was operated at was 25kW. Additional testing was conducted using a NASA

cathode designed for higher current operation. During this testing, thrust over 1 N was measured at 40.2 A and 548.9 V.

Several issues related to operation of Hall thrusters at these high powers were encountered.

14. SUBJECT TERMS

Hall thruster

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION

OF REPORT OF THIS PAGE

Unclassified Unclassified

NSN 7540-01-280-5500

19. SECURITY CLASSIFICATION

OF ABSTRACT

Unclassified

115. NUMBER OF PAGES

19

16. PRICE CODE

A03

20. LIMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18

298-102