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AIAA-98-3792

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Test Results of a 200W Class Hall Thruster

David T. Jacobson

Dynacs Engineering Company, Inc., Brook Park, Ohio

Robert S. Jankovsky

Glenn Research Center, Cleveland, Ohio

December 1999

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NASA / TM--1999-209449 AIAA-98-3792

Test Results of a 200W Class Hall Thruster

David T. Jacobson

Dynacs Engineering Company, Inc., Brook Park, Ohio

Robert S. Jankovsky

Glenn Research Center, Cleveland, Ohio

Prepared for the

Joint Propulsion Conference

sponsored by the American Institute of Aeronautics and Astronautics

Cleveland, Ohio, July 13-15, 1998

National Aeronautics and

Space Administration

Glenn Research Center

December 1999
Acknowledgments

The authors would like to thank Mr. Michael Patterson of NASA GRC for the use of the laboratory cathode and his

support in this effort. This piece of hardware was necessary to demonstrate the capabilities of the low power

Hall thruster system. The authors would also like to thank Mr. Luis Pinero of NASA GRC and Glen Bowers

of Gilcrest Electric for the design and fabrication of the BBPPU. A special thanks is extended to Patrick Barber

of NASA's Test Installation Division for fabrication of the experimental hardware. His hard work and

dedication is greatly appreciated.

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
TEST RESULTS OF A 200W CLASS HALL THRUSTER

David T. Jacobson

Dynacs Engineering

Brook Park. Ohio 44142

Robert S. Jankovsky

National Aeronautics and Space Administration

Gllenn Research Center

Cleveland, Ohio 44135

ABSTRACT

The performance of a 200 W class Hall thruster was evaluated. Performance measurements were taken at

power levels between 90 W and 250 IV. At the nominal 200 W design point, the measured thrust was 11.3

raN, and the specific impulse was 1170 s excluding cathode flow in the calculation. A laboratotx model 3

mm diameter hollow cathode was used for all testing. The engine was operated on laboratory power

supplies in addition to a breadboard power processing unit fabricated from commercially available DC to

DC com'erters.

INTRODUCTION

Over 70 Stationary Plasma Thrusters (SPT's)

have been flown on Russian spacecraft since

1971.' Recently, there has been increasing U.S.

Government and commercial interest in utilizing

this technology on Western spacecraft due to the

tremendous economic benefits and the

maturation of Hall thruster systems. The

Ballistic Missile Defense Organization (BMDO),

through NASA. was at the forefront of efforts to

demonstrate the benefits of Hall thruster

technology with extensive testing of the SPT100 beginning in 1993.: Since that time there

has been an on-going research effort to evaluate

Hall thruster technology.

NASA has recently placed an emphasis on

smaller, lower cost, lower power spacecraft for

space science missions. Propulsion system

options for these small spacecraft are currently

limited. Recent mission analysis indicated that a

low-power Hall thruster system has mission

enabling advantages over small spacecraft

equipped with state-of-the-art (SOA) chemical

systems.' In an effort to demonstrate the

feasibility of a very low-power Hall system,

NASA, through Atlantic Research Corporation,

obtained a 200 W class Hall thruster built by the

Research Institute of Applied Mechanics and

Electrodynamics of Moscow Aviation Institute

(RIAME). A performance evaluation of this

engine was conducted and is presented in this

report.

The main objective of this evaluation was to

demonstrate the capabilities of very low-power

laboratory model Hall thrusters. The second

objective was to demonstrate the application of a

low power 3 mm diameter hollow cathode being

developed by NASA. The final objective was to

illustrate the potential of building a low cost. low

power Hall thruster system. The main activity

associated with this objective was the design and

fabrication of a breadboard power processing

unit (BBPPU).

APPARATUSANDPROCEDURE

The Hall thruster used in this investigation was

an SPT-30 obtained from the Research Institute

of Applied Mechanics and Electrodynamics of

Moscow Aviation Institute (RIAME). The

engine tested was a laboratory model thruster

which would require additional development to

achieve flight ready status. The SPT-30 derives

its name from the outer diameter of the discharge

chamber which is 30 ram. It utilizes separate

coaxially wound inner and outer magnets,

isolated from the discharge and each other, to

generate the required magnetic field. The engine

was designed to operate in the power range of

100-200 W and has an estimated lifetime

NASA/TM--1999-209449 1
exceeding 600 hours. Performance testing of the

engine was conducted by RIAME prior to this

investigation. A photograph of the thruster

installed on a thrust balance along with a 3 mm

NASA hollow cathode is shown in Figure 1.

A 3 mm laboratory cathode based on the NASA

Plasma Contactor design originally developed

for the International Space Station (ISS) was

used to operate the engine for the performance

testing. _ Figure 2 is a photograph of this

cathode. The cathode utilizes a low work

function insert for electron emission. A

resistance heater is used for pre-operation

conditioning after exposure to atmosphere, and

heating prior to ignition. The cathode was

designed to operate with emission current

between 0.5-1.5 A with xenon flow rates greater

than 0.I mg/s.

The thrust stand used for this test sequence was

of an inverted pendulum type used previously

for Hall thruster performance measurements. _

Recent modifications to the thrust stand

electronic controls permitted thrust

measurements to be performed using null rather

than displacement method. The null type thrust

stand is composed of three components

including an electromagnetic driver,

proportional, integral, differential, (PID)

controller, and a Linear Variable Displacement

Transformer (LVDT). The null method of

measurement relies on current flow in a coil to

generate an acceptable force via a magnetic field

required to maintain a zero displacement. The

current was measured using a shunt when the

thruster was operating and output recorded using

a strip chart recorder. Thrust calibrations were

performed before, during and after testing.

Three 0.48 g weights were applied using a

stepping motor, providing a known coil current

for a known force applied. Using this calibration

data. the actual thrust produced by the engine

was calculated. Reported thrust values have an

estimated uncertainty of + 2.0%.

Commercially available power supplies were

used to operate the thruster and cathode for all

performance measurements. Separate power

supplies were used for the discharge, inner and

outer magnets, cathode heater and keeper. A low

voltage, high current keeper supply protected by

a blocking diode was used in parallel with a high

voltage, low current ignitor supply for cathode

ignition and operation. An output filter

consisting of a 1.0 f2 ballast resistor in series

with the anode and 100 IaF capacitor between

cathode and anode was used. Back to back 50

volt Zener diodes were used to protect the power

supplies from an anomalous operating condition.

An electrical schematic diagram of the power

supply configuration is shown in Figure 3.

The engine was also successfully operated on a

breadboard power processing unit (BBPPU)

fabricated internally at NASA Glenn Research

Center (GRC) from commercially available DC

to DC converters. These commercial power

modules operated from an input voltage of 21 to

32 Vdc provided regulated voltage, up to 285

Vdc, for the discharge and regulated currents for

the inner and outer magnets and the cathode

keeper and heater. These outputs can be

individually adjusted through a wide range to

optimize thruster operation. The power modules

demonstrated an efficiency of 0.80 to 0.90, high

power density, and functions including

overcurrent and overvoltage protection.

Xenon flow was provided to the cathode and

anode by a laboratory feed system which

incorporated commercially available, National

Institute of Standards and Technology (NIST)

traceable flow controllers. Flow calibrations

were performed before and after pertormance

data was taken using a constant volume method.

The uncertainty associated with the mass flow

was estimated to be + 1.0%. Commercially

available research grade xenon with purity

99.999% was used. The feed system was helium

leak checked prior to operation of the

propulsion system.

All performance testing was conducted in the

main volume of Vacuum Facility 8 at NASA

GRC. The 1.5 m diameter and 5 m long

chamber was equipped with a pumping train

consisting of four 0.89 m oil diffusion pumps

backed by two roughing pumps and a

mechanical blower. The facility was capable of

maintaining a background pressure of 1.0 x 10_

Torr, measured with ion gauges calibrated on air

and uncorrected for xenon, with the thruster

operating at the highest mass flow rate.

NASA/TM--1999-209449 2
Thrusterperformancewasinvestigatedoverthe

powerrange90to260W. Thrusteroperationat

apowerlevelofgreaterthan200Wwaslimited

to a periodof 5 rain as suggestedby the

manufacturer.Dischargecurrentandvoltage,

cathodeto groundvoltage,andinnerandouter

magnetvoltagesweremeasuredusingcalibrated

digitalmultimeters.Thedischargecurrentand

voltagewerealsorecordedat eachdatapoint

using a digital oscilloscopewith storage

capability.Datawascollectedatafixedanode

massflowratewhilevaryingdischargevoltages

between100and250Vin50Vincrements.

Thecathodewaspreheatedpriortoignitionat9

A of heatercurrentwitha massflowrate0.5

mg/s. With thekeepersupplysetat 60 V,

breakdowntypicallyoccurredwithin2 minutes.

Theignitorsupplywasonlyusedonthefirsttwo

starts,with ignitionvoltageof approximately

250 V. Successfulcathodeignition was

performedat lowermassflowratesof 0.3and

0.2mg/slaterinthetestsequence.

Thedischargewasinitiatedby turningonthe

dischargepowersupplywiththecurrentlimitset

at1Aandthecathodeoperatingonthekeeperat

0.5A withaflowrateof 0.1mg/s.Theinner

andoutermagnetcurrentwassetat 1Aandthe

dischargevoltagewas rampedup to until

breakdownoccurred. The main discharge

typicallystartedatvoltagesbetween35and50

V. Followinginitiationof thedischarge, the

dischargevoltagewasincreasedto thedesired

value,whileincrementallyincreasingtheinner

andoutermagnetcurrentvaluesto maintain

stableoperation. The startingprocedure,

implementedwithoutmodification, includingthe

magnetcurrentvalues, was suggested by

RIAME. Because the thruster was stored for an

extended period of time in an uncontrolled

environment following delivery, the thruster was

operated at low voltage in excess of one hour the

first time the engine was operated. This allowed

for outgassing of the hydroscopic insulators at an

acceptably low rate.

RESULTS AND DISCUSSION

During this investigation, performance

measurements were taken at power levels from

90 to 260 W. Data was collected at discharge

voltages ranging from 100 to 250 V and

discharge currents corresponding to anode mass

flow rates of 0.98, 0.75 and 0.61 mg/s. Data was

taken at a nominal cathode mass flow rate of 0.1

mg/s. Although no single design point was

specified by the manufacturer, discharge voltage

and current values of 200 V and 1.0 A produced

a power consistent with the maximum design

power level. Performance values at a nominal

operating point of 200 W indicate specific

impulse of 1170 s with an efficiency of 0.32

excluding cathode flow. and a specific impulse

of 1050 s with an efficiency of 0.29 including

cathode flow. Specific impulses as high as 1290

s and efficiencies up to 0.29 were measured at

the same power level, but at a correspondingly

lower thrust. A complete table of the

performance test results is provided as Table 1.

The relationship between the thrust produced by

the SPT-30 and discharge power at three anode

floss, rates is shown in Figure 4. Because this

particular engine was not designed with a

magnetic system optimized with respect to

power, power was calculated using only the

discharge voltage and current and does not take

into account the power dissipated in the magnets.

The measured thrust produced was between 4.9

and 13.2 mN. Both the performance data taken

in Vacuum Facility 8 at NASA GRC and that

taken at the RIAME facility are included in this

plot. Open symbols indicate data taken at NASA

GRC and closed symbols indicating data taken at

the RIAME facility. Two of the three data sets

were taken at essentially the same anode mass

floss,, while the maximum flow rate at which data

was taken was slightly higher in the test

sequence performed at NASA GRC. All data

series exhibit similar trends with the higher

thrust produced at the highest mass flow rate.

Scatter among points at essentially the same

power can be attributed to differences in

discharge current as a result of non-optimized

magnet settings, and differences in discharge

voltage values.

Specific impulse as a function of thruster power

is shown in Figure 5 for three anode flow rates.

Closed symbols indicate values calculated

excluding cathode flow and open symbols

cathode including cathode flow. Data were

plotted for calculations performed with and

without cathode floss' to illustrate the effect of

cathode flow on performance. This effect arises

because propellant expended through the

cathode is not accelerated and does not

NASA/TM--1999-209449 3
contributetothethrustproducedbytheengine.

Thehigherspecificimpulsevaluesoccurredat

thelowestanodemassflow rates,andin all

cases,highervalueswererecordedat higher

dischargevohages.A greaterimpactof cathode

flowonperformanceisapparent aloweranode

massflowratesbecausecathodeflowisagreater

fractionof thetotalmassflow.Thecathodeflow

fractionforthissystemrangedfrom0.II to0.16.

Furtherminiturationof the cathodewould

significantlyimprovetheperformanceof small

Hallthrusters.

Thrusterefficiencyas a functionof thruster

powerisshowninFigure6forthreeanodemass

flowrates.Onceagain,datawereplottedboth

withandwithoutcathodeflowandthesymbol

conventionremainsconsistent.Theefficiency

valuescalculatedincludingcathodeflowwere

only3-5%lessthanthosecalculatedexcluding

cathodeflow. Thiseffectof cathodeflowon

performancewas significantlyreducedin

comparisontodatacollectedpreviouslyonother

similarlow powerHallsystemswhichdid not

utilizealowflowcathodedesign.:

Thecurrentandvoltagecharacteristics of the

enginewererecordedateachdatapoint.Typical

valuesof peakto peakdischargecurrent

oscillationsrangedfrom 0.2 to 1.1A. The

currentoscillationsfromthe dischargepower

supplywereminimal.Voltageoscillationswere

alsominimalateachpoint.

In all cases,the procedureoutlinedin the

previoussectionwasusedtostartthedischarge

withoutanyanomalousbehavior,butdifficulty

sustaininga dischargewhile adjustingthe

magnetcurrentwasfrequentlyexperienced. As

aresult,themethodof magnetoptimizationby

minimizationofthedischargecurrentforagiven

dischargevoltageand flow rate was not

implemented. _ Instead,two magnetcurrent

setpointswere identifiedat which thruster

stabilitywasacceptable, andperformancedata

weretakenat thosepoints.It is suspectedthat

thisbehaviorwaspartiallya resultof emissive

characteristics of thecathode.A modification

wasmadeto thekeeperorificeto increasethe

transparency oftheplasmawheninitialdifficulty

wasencounteredstartingthe engine. Also,

supplyinga smallamountof keepercurrent

aidedin maintaininga discharge: although,no

performancedatawasrecordedwiththekeeper

on. Engineoperationwiththecathodeheater

operatingwill be performedin future

investigations to eliminatethepossibilityof the

cathodeextinguishingitself as a resultof

temperature.

Thecathodeorientationthatwasimplemented

wasa resultof initial difficultyencountered

startingtheengine.Aninitialattemptwasmade

tooperatetheenginewiththecathodepositioned

inamoretraditional locationwiththeaxisofthe

cathodeat a45° angleto theexitplaneof the

thruster. The configurationof the keeper

imposedgeometricconstraintswhichestablished

thedistancethatthecathodewaslocatedfrom

the dischargechamber. Thecenterof the

cathodeorificewas20mmfromtheexitplane

of thethrusterandthe20mmfromtheouter

diameterof thedischargechamber.In this

configuration,severalunsuccessful attempts

weremadetoinitiateadischarge.Thecathode

was re-orientedsuchthat the cathodewas

perpendicular tothethrustaxis,andthedistance

fromtheorificeto thedischargechamberwas

minimized. No startingdifficultieswere

encounteredwith the cathode in this

configuration, andall performancetestingwas

conductedwith the cathodeorientedin this

fashion.Cathodepositionseemsto havesome

impacton facilitatingthe initiationof the

discharge.Furtherinvestigationwouldneedto

beconductedtodeterminethemagnitudeof the

effect of cathodeposition on starting

characteristicsandperformanceto makeany,

definitiveconclusion.

CONCLUDING REMARKS

An investigation was conducted to evaluate the

feasibility of using Hall thruster technology for

very low power applications, using a low power

laboratory model 200 W class Hall thruster. Data

were collected at discharge voltages between

100 and 250 V without exceeding a power level

of 200 W for a periods of more than 5 min. The

thruster was capable of producing 11.3 mN of

thrust at a nominal power level of 200 W. The

specific impulse and efficiency values

corresponding to that point were 1170 s and

0.32, respectively, excluding cathode flow.

All testing was conducted using a NASA

laboratory model, low flow, 3 mm diameter

NA SA/TM-- 1999- 209449 4
hollowcathode.Successful operationof the

SPT-30wasdemonstrated for emissioncurrent

between0.6and1.0A. Insomecasesdifficulty

wasencounteredsustainingenginedischarge.

Thismayhavebeenrelatedtocathodeposition

ortemperature. Thelaboratoryunitusedforthis

pertbrmancetestwasnotoptimizedwithregard

to keepergeometryor operationof a Hall

thruster.Operationofthekeeperorheaterwhile

performancemeasurements werebeingtaken

mayhaveaidedin sustainingadischarge. The

benefitof a low flow hollow cathodeon

performancewasdemonstrated.

Operationof theengineonabreadboardpower

processingunitwasdemonstrated. Todate,the

BBPPUwasusedtooperatethedischargeupto

voltagesof 275 volts. Modificationsare

currentlybeingimplementedto allow for

operationof the entiresystemwithoutthe

combineduseoflaboratorypowersupplies.

REFERENCES

Clauss. C.W.. Tilley, D.L. and Barnhart,

D.A., "Benefits of Low-Power Stationary

Plasma Thruster Propulsion For Small

Satellites," Proceedings of the 9th Annual

AIAA/USU Conference on Small Satellites,

Utah, Sept. 1995

2. Sankovic. J.M., Hamley, J.A., and Haag,

T.W., "Performance Evaluation of the

Russian SPT- 100 Thruster at NASA LeRC,'"

IEPC-93-094, Sept. 1993.

3. Oleson, S., "'Electric Propulsion For Low

Earth Orbit Communication Satellites."

IEPC-97-148, August, 1997.

4. 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

tbr Space Sation Freedom," AIAA-93-2228,

June 1993.

5. Hagg, T.W., and Osborn. M..

"RHETT/EPDM Performance

Characterization," IEPC-97-107, August

1997.

6. Sankovic, J.M.. Haag, T.W., and Manzella,

D.H., "'Operating Characteristics of the

Russian D-55 Thruster wire Anode Layer."

AIAA-94-3011. June 1994.

7. Manzella, D.H,, Oleson. S., Sankovic. J.M.,

Haag, T.W., Semenkin, A., and Kim, V.,

"Evaluation of Low Power Hall Thruster

Propulsion," AIAA-96-2736, July, 1996.

NASA/TM--1999-209449 5
Tablel.PerformancevaluescollectedfortestingofSPT-30Hallthruster.

DischargDischargeAnodeCathodePower Thrust SpecificEfficiency Specific Efficiency*

e Current Flow Flow Impulse Impulse*

Voltage

V A mg/s mg/s W naN s s

200 1.01 0.98 0.11 202.0 l 1.3 1050 0.29 1170 0.32

150 1.02 0.98 0.11 153.0 8.8 823 0.23 914 0.26

100 0.99 0.98 0,10 99.0 5.6 521 0.14 576 0.16

250 1.03 0.98 0.11 257.5 13.2 1230 0.31 1370 0.34

150 1.02 0.98 0.11 153.0 8.8 822 0.23 915 0.26

150 0.79 0.75 0.10 118.5 6.9 818 0.23 931 0.27

200 0.78 0.75 0.10 156.0 8.3 986 0.26 1120 0.29

250 0.82 0.75 0.10 205.0 9.5 1130 0.26 1290 0.29

150 0.75 0.75 0.10 112.5 6.5 775 0.22 882 0.25

150 0.60 0.61 0.10 90.0 4.9 700 0.19 819 0.22

200 0.60 0.61 0.10 120.0 6.2 874 0.22 1020 0.26

250 0.63 0.61 0.10 156.3 7.0 1000 0.22 1170 0.26

* denotesefficiencyandspecificimpulsevaluescalculatedexcludingcathodeflow

Figure1.SPT-30onthruststandinvacuumfacility8atNASAGlennResearchCenter.

NASA/TM--1999-209449 6
Figure 2. NASA Laboratory Model. low-flow cathode.

R = 1 £2 Anode

Discharge Supply [

_+ Heater

Igm'tor

C = I OOJ3F:

Ca thode

Thruster

__k__

t"

I

Inner Magnet Supgly

IInner Magnet +

Inner Magnet -

Outer Magnet + "t

I

OuterSulslAy Magnet [

i- Outer Magnet -

Figure 3. Electrical Schematic diagram.

NASAFFM--1999-209449 7
14

12

10

8

6

4

2

0

0

Figure 4.

O

O•

O

8

[] m

0 0.98 mg/s

ri 0.75 mg/s

A0.61 mg/s

• 0.88 mg/s

• 0.76 mg/s

• 0.60 mg/s

I I I I I

50 IO0 15O 200

Thruster Power (W)

250 300

Open Symbol-NASA GRC Data

Closed SymboI-RIAME Data

Comparison of performance data taken at NASA GRC and RIAME.

1400

1200

1000

800

600

4O0

200

0

Figure 5.

A

O

ii 'h

• e,

•A

[] o

[]

o

o

0 0.98 mg/s

[] 0.75 mg/s

/_0.61 mg/s

• 0.98 mg/s

• 0.75 mg/s

• 0.61 mg/s

I I I I I

50 100 150 200 250 300

Thruster Power (W)

Open Symbols-including cathode flow

Closed Symbol-excluding cathode flow

Specific Impulse verse thruster power both with and without cathode flow.

NASA/TM-- 1999-,.09449 ,.,} 8
0.35

0.3

0.25

0.2

0.15

0.1

0.05

A

•|

[]

rnA

O

I

O

A

fll

[]

O

O 0.98 mg/s

[] 0.75 mg/s

A0.61 mg/s

• 0.98 mg/s

• 0.75 mg/s

• 0.61 mg/s

I I I I I

0 50 100 150 200 250 300

Thruster Power (W)

Open Symbols-including cathode flow

Closed Symbol-excluding cathode flow

Figure 6. Efficiency verse thruster power both with and without cathode flow.

NASA/TM--1999-209449 9
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4. TITLE AND SUBTITLE

Test Results of a 200W Class 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

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National Aeronautics and Space Administration

Washington, DC 20546-0001

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NASA TM--1999-209449

AIAA-98-3792

11. SUPPLEMENTARY NOTES

Prepared for the Joint Propulsion Conference sponsored by the American Institute of Aeronautics and Astronautics.

Cleveland. Ohio. July 13-15, 1998. David T. Jacobson, Dynacs Engineering Company, Inc., Brook Park, Ohio 44142,

and Robert S. Jankovsky, NASA Glenn Research Center. Responsible person, Robert S. Jankovsky, organization code

5430. (216) 977-7515.

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13, ABSTRACT (Maximum 200 words)

The performance of a 200 W class Hall thruster was evaluated. Performance measurements were taken at power levels

between 90 W and 250 W. At the nominal 200 W design point, the measured thrust was 11.3 mN, and the specific impulse

was I 170 s excluding cathode flow in the calculation. A laboratory model 3 mm diameter hollow cathode was used for all

testing. The engine was operated on laboratory power supplies in addition to a breadboard power processing unit fabricated from commercially available DC to DC converters.

14. SUBJECT TERMS

Hall thruster

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A03

20, LIMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18

298-102