NASA/TM--2002-211887 AIAA-2002-3676

Laboratory Model 50 kW Hall Thruster

David Manzella

University of Toledo, Toledo, Ohio

Robert Jankovsky

Glenn Research Center, Cleveland, Ohio

Richard Hofer

QSS Group, Inc., Cleveland, Ohio

September 2002
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NASA/TM--2002-211887 AIAA-2002-3676

Laboratory Model 50 kW Hall Thruster

David Manzella

University of Toledo, Toledo, Ohio

Robert Jankovsky

Glenn Research Center, Cleveland, Ohio

Richard Hofer

QSS Group, Inc., Cleveland, Ohio

Prepared for the

38th Joint Propulsion Conference and Exhibit

cosponsored by the AIAA, ASME, SAE, and ASEE

Indianapolis, Indiana, July 7-10, 2002

National Aeronautics and

Space Administration

Glenn Research Center

September 2002
NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

Available from

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

Available electronically at _sa. _ov
Laboratory Model 50 kW Hall Thruster

David Manzella

University of Toledo

Toledo, Ohio 43606

Robert Jankovsky

NASA Glenn Research Center

Cleveland, Ohio 44135

Richard Hofer

QSS Group, Inc.

Cleveland, Ohio 44135

A 0.46 meter diameter Hall thruster was fabricated andperformance tested atpowers up to 72 kilowatts. Thrusts up

to 2.9 Newtons were measured. Discharge specific impulses ranged from 1750 to 3250 seconds with discharge

efficiencies between 46 and 65 percent. Overall specific impulses ranged from 1550 to 3050 seconds with overall

efficiencies between 40 and 57 percent. Performance data indicated significant fraction of multiple-charged ions

during operation at elevated power levels. Cathode mass flow rate was shown to be a significant parameter with

regard to thruster efficiency.

Introduction

High power electric propulsion systems have been

shown to enable a number of missions including

missions to Mars and Earth orbital solar electric

power generation for terrestrial use. 1'2 These types of

missions require moderate transfer times and sizable

thrust levels resulting in an optimized propulsion

system specific impulse from 2000-3000 seconds

based on the available on-board power. Hall thruster

technology offers a favorable combination of

performance, reliability, and lifetime for such

applications based on the characteristics of state-ofthe-art systems. As a result, the NASA Space Solar

Power Concept and Technology Maturation Program

initiated preliminary strategic technology research

and development into high power Hall thruster

technology to enable space solar power systems and

other high power spacecraft.

High power Hall thruster technology for primary

propulsion applications was initially investigated in

the former Soviet Union and later Russia for

interplanetary missions. 3'4 Efforts conducted at the

FAKEL Design Bureau in Kaliningrad, Russia

culminated in the SPT-290. This thruster utilized a

290 mm outer diameter ceramic channel to ultimately

produce as much as 1.1 Newtons of thrust at

25 kilowatts of input power using xenon as the

propellant. Efforts at the Central Scientific Institute

for Machine Building (TsNIIMASH) in Korolev,

Russia resulted in the demonstration of a

100 kilowatt two-stage anode layer thruster

demonstrating a specific impulse of 8000 seconds

with a discharge efficiency of 80 percent using

bismuth as the propellant. Later a xenon fueled anode

layer thruster was tested at power levels up to

25 kilowatts demonstrating performance similar to

the SPT-290 at that power level. 5 More recently an

SPT-type thruster designated the T-220, which was

developed in the United States under contract to

NASA Glenn Research Center by TRW in

cooperation with Space Power Incorporated (who has

since become a part of Pratt & Whitney's Space

Propulsion and Chemical Systems Division), was

tested. 6 This thruster which has a 220 mm outer

diameter ceramic channel was originally designed to

produce 0.5 Newtons of thrust at 10 kW but has been

modified to produce in excess of 1 Newton of thrust

at elevated power levels]

Development of a Hall thruster capable of operating

at 2000-3000 seconds at power levels of 50 kilowatts

and above requires a thruster design which preserves

the physical processes required for efficient

ionization and acceleration of the propellant.

Important characteristics for preservation of these

processes include: discharge current density,

discharge chamber geometry, and magnetic field

distribution. 8 This paper describes the development

of a Hall thruster designed to operate at 50 kilowatts.

Test results including the effect of discharge voltage,

anode mass flow rate and cathode mass flow rate on

thruster performance are included.

Apparatus

A photograph of the NASA-457M 50-kilowatt

laboratory Hall thruster is shown in figure 1. The

NASA/TM--2002-211887 1
thruster, whichhasa457mmouterdiameter ceramic

dischargechamber, injectsxenonpropellant intothe

rearof thechannel throughaseriesof holeslocated

alongtheinnerandouterwallsof theannular

metallicanodelocatedattherearofthechannel.The

cathode, whichhasbeenpreviouslydescribed, 9was

locatedonthethrustercenterline.

A magneticfieldisestablishedacrosstheexitplane

of thedischargechamberthroughtheuseof two

concentricelectromagnets. Eachelectromagnet can

beindependently powered. Theelectromagnets along

withmagneticpolesandscreensformthemagnetic

circuit.Thedesignof themagneticircuitwas

optimizedthroughtheuseof acommercial threedimensional magneto-static computer code,Magnet6

byInfolytica.Thefieldlinetopography employedin

theNASA457Mis qualitativelysimilarto the

plasmalensdesignoftheNASA173M. 1°Following

fabricationof the thrustertheresultsfromthe

magneticmodelwerecomparedwith magnetic

intensitiesmeasuredwitha3-axisGaussmeter.The

measuredandcalculatedradialandaxialcomponents

of themagneticfieldareshownin figure2. The

channel depthnormalizedtheaxialdistanceandboth

theradialandaxialmagneticfieldcomponents were

normalizedby themaximumvaluesrespectively.

Theinneradialmagneticfieldstrengthwasslightly

higherthancalculatedwhichalsoaffectedthe

centerlineradialfieldstrength. Thisdifferencewas

attributedto uncertaintyin thepropertiesof the

material usedinthispartofthemagneticcircuit.

Thethrusterwastestedin a cryogenicallypumped

cylindrical vacuumchamber measuring5 metersin

diameter by20metersinlength.Thisfacility,which

hasbeendescribedin detailpreviously, l° was

modifiedfor this test.A cylindricaltestport

measuring2 metersin diameterby 2.5metersin

lengthwasmountedalongthemajoraxisofthetank

witha2-meterisolationvalvebetweenthetestport

andthemainvolumeofthetank.Aphotographofthe

testfacilityis shownin figure3. Theeffective

pumpingspeedof the test facility in this

configurationwas700,000literspersecond. Thiswas

determinedbasedonthexenonmassflowrateand

thetankpressuremeasuredbehindthethruster.Asa

result,thehighestbackgroundpressurewas2x 10.5

Torrataxenonflowrateofjustover1standardliter

perminute(SLPM).Thispressureisstillbelowthe

maximumrecommendedvaluefor performance

testingof 5x 10.5Torr. 11 A laboratory xenon feed

system, required auxiliary power supplies, and a data

acquisition system were located adjacent to the test

port on an elevated deck platform that was also

modified for these tests. The discharge power supply

was a remotely located three phase commercially

available unit with a rated output of 100 Amperes and

2000 Volts. An output filter consisting of a 21mF

capacitor located between the anode and cathode was

used. The thrust produced by the NASA-457M was

measured using an inverted pendulum design thrust

stand fabricated specifically for these tests. The thrust

stand was designed based on a thrust stand used in

previous evaluations of Hall effect thrusters. 12'13

Conceptually the thrust balance was identical to the

previous design except for modifications necessary to

accommodate the increased mass and currents. The

thrust stand can be seen beneath the thruster in the

photograph in figure 4. Due to the large thermal

mass of the NASA-457M thruster no attempts were

made to establish complete thermal equilibrium prior

to measuring performance data. The thruster was

typically operated for approximately an hour to allow

not only the thruster discharge chamber to warm up,

but to also allow the thrust stand to equilibrate with

the operating thruster. Following this warm up

period performance data was taken. Thrust stand

calibrations, conducted by applying a series of one

hundred gram weights, were conducted before and

after each series of test data.

Results and Discussion

The NASA-457M was tested over a range of input

powers from 9 to 72 kilowatts by varying the anode

mass flow rate from 15 to 93 mg/s and the discharge

voltage from 300 to 650 Volts. All these data are

presented in an appendix at the end of this paper.

Over this range of operating conditions thruster

operation was stable. Thrust is plotted as a function

of discharge power for five different discharge

voltages in figure 5. For each voltage the maximum

current corresponded to approximately 110 Amperes.

This maximum resulted from the maximum current

capacity of the discharge power supply. As can be

seen from the figure, thrust varied linearly with

discharge power for a given discharge voltage. At

650 Volts a thrust of 2.9 Newtons was measured at

111 Amperes. There was no evidence of thermal

limitations at this power level although it was noted

that the anode was glowing upon shutdown. While

this has been observed with other thrusters operating

at nominal operating conditions at shutdown, this was

not observed at powers up to 50 kW with the NASA457M. Operation at voltages above 650 Volts was

investigated but was not adequately characterized due

to voltage isolation problems within the thruster.

Modification of the isolation scheme is currently

underway which should allow for operation up to at

1000 Volts for future testing.

The reduction in thrust with increasing voltage at a

given power level was accompanied by a

corresponding increase in specific impulse. This is

NASA/TM--2002-211887 2
showninfigure6wherespecificimpulseisplotted

versusdischargepower.Dischargespecificimpulse

rangedfrom approximately1750secondsto

3250seconds. Theoverallthrusterspecificimpulse

includingthecathodemassflowrateandmagnet

powerangedfrom1550secondsto 3050seconds.

Whileoverallspecificimpulseandefficiencyare

reported, it shouldbenotedthatnoattemptswere

madeto optimizethelaboratorymodelcathodeor

magneticcircuit for efficientoperation.The

dischargespecificimpulseat a givenvoltage

comparesfavorablywith other state-of-the-art

thrusters. Forexampleat300VoltsanSPT-100hasa

dischargespecificimpulseof 1750seconds 14and

thecomparablysizedhighvoltageSPT-1thruster

providedbetween2250and 2500secondsat

500Volts. 15 The functional dependence of specific

impulse with discharge voltage was considered in

more detail by Hofer where experimental data were

compared with various predicted values. 16 These

predictions suggest as discharge power increased

above 20 kilowatts or the discharge current increased

above 30 Amperes that the effect of multiple-charged

ions becomes significant.

The variation of discharge current versus anode mass

flow rate is shown in figure 7. The dependence is

linear for anode mass flow rates below approximately

70 mg/s. Above this mass flow rate the discharge

current increased more rapidly with increasing anode

mass flow rate. This increase in current at high

anode mass flow rate was the result of either an

increase in the electron or ion current contribution to

the discharge current. The corresponding specific

impulse and higher density within the channel

suggest that this effect is due to additional ion current

from multiple-charged ions.

The effect of multiple charged ions on specific

impulse was previously considered when Hofer

predicted 100 Amperes at 88.8 mg/s and 500 Volts

with a 2768 second specific impulse for a multiplecharged plasma. For a singly charged plasma

100 Amperes at 101 mg/s and 500 Volts with a

2622 second specific impulse was predicted. The

measured anode mass flow rate at 500 Volts and

100 Amperes was 86.4 mg/s at just less than

2750 seconds, which agrees favorably with the

multiple-charged prediction. This prediction and

other predicted values of specific impulse are shown

in figure 8 versus discharge voltage for an anode

mass flow rate of 86.4 mg/s. Experimental data are

also included.

During the course of this investigation the effect of

cathode flow rate on thruster performance was also

considered. Previously the sensitivity of cathode

mass flow rate on thruster operation was considered

in detail for a 4.5 kW Hall thruster. 17 As was pointed

out in this prior investigation, the coupling of the

plasma produced in the hollow cathode with the

plasma produced

channel is not

complicated than

simple external

magnetic fields.

in the Hall thruster discharge

fully understood and is more

running a hollow cathode with a

anode due to the presence of

Figure 9 shows the cathode-toground voltage as a function of discharge current for

three different cathode flow rates. As can be seen

from the figure, for each cathode flow rate the

cathode-to-ground voltage varied linearly with

discharge current with the voltage becoming more

negative with increasing current.

For comparison purposes, this cathode was tested

independent of the Hall thruster with an external

anode at a mass flow rate of 2 mg/s. 9 For currents

between 50 and 100 Amperes the coupling voltage

was between -10 and -13 volts. This suggests that

the additional voltage required to couple the plasma

from the cathode to the plasma from the Hall thruster

was associated with electron transport across the

applied magnetic field. As such, this additional

voltage represents a thruster loss mechanism since a

larger portion of the voltage applied between the

cathode and anode is needed for cathode coupling

leaving a smaller portion of the total applied voltage

for ion acceleration.

The effect of large cathode-to-ground coupling

voltages on thruster discharge efficiency is shown in

figure 10. These data show that the smaller the

cathode-to-ground voltage the higher the discharge

efficiency (and overall efficiency as can be seen in

the data tables). This substantiates the conclusion that

large cathode-to-ground voltages were indicative

of an energy loss mechanism associated with

poor cathode coupling. The data also indicate a

peak discharge efficiency of near 65 percent at

around 500 Volts. The efficiency decreases with

increasing voltage above this value. This is consistent

with past investigations into high voltage Hall

thruster operation. 15'is

Conclusions

A 0.46 m outer diameter Hall thruster was fabricated

and performance tested at powers up to 72 kilowatts.

These tests demonstrated the efficacy of scaling Hall

thrusters to high power suitable for a range of future

missions. Thrusts up to nearly 3 Newtons were

measured. Discharge specific impulses ranged from

1750 to 3250 seconds with discharge efficiencies

between 46 and 65 percent. Overall specific impulses

ranged from 1550 to 3050 seconds with overall

efficiencies between 40 and 57 percent.

NASA/TM--2002-211887 3
Performance data suggested a significant fraction of

multiple-charged ions during operation at elevated

power levels. This conclusion was supported by

previous performance predictions by Hofer and the

functional dependence of discharge current with

anode mass flow rate.

An investigation into the effect of cathode flow rate

on thruster operation was conducted. Cathode mass

flow rate was shown to be a significant parameter

with regard to thruster efficiency. It was also

demonstrated that the coupling voltage between the

cathode and thruster anode was significantly different

than that measured with the cathode operating to a

planar external anode.

References

1. Gefert, L., Hack, K., and Kerslake, T., "Options

for the Human Exploration of Mars Using Solar

Electric Propulsion," 1999.

2. Oleson, S., "Advanced Propulsion for Space

Solar Power Satellites," AIAA-99-2872, June

1999.

3. Loeb, H. and Popov, G., "Advanced

Interplanetary Mission of the XXI Century Using

Electric Propulsion," IEPC-95-04, Sept. 1995.

4. Garkusha, V., et al., "Electric Propulsion

Activity in TsNIIMASH," IEPC-95-09, Sept.

1995.

5. Jacobson, D. and Jankovsky, R., "Performance

Evaluation of a 50 kW Hall Thruster," AIAA99-0457, January 1999.

6. Jankovsky, R., McLean, C., and McVey, J.,

"Preliminary Evaluation of a 10kW Hall

Thruster," AIAA-99-0456, January 1999.

7. Britt, N., "Electric Propulsion Activities in U.S.

Industry," AIAA-2002-3559, July 2002.

8. Arkhipov, B., et al., "Development and

Investigation of Characteristics of Increased

Power SPT Models," IEPC-93-222, Sept. 1993.

9. Carpenter, C. and Patterson, M., "High-Current

Hollow Cathode Development," IEPC-01-274,

October 2001.

10. Hofer, R. R., Peterson, P. Y., Gallimore, A. D.,

"A High Specific Impulse Two-Stage Hall

Thruster with Plasma Lens Focusing," IEPC-01-

036, 27 th International Electric Propulsion

Conference, Pasadena, CA, Oct 14-19, 2001.

10. Grisnik, S., and Parkes, J., "A Large High

Vacuum, High Pumping Speed Space Simulation

Chamber for Electric Propulsion," IEPC93-151, Sept. 1993.

11. Randolph, T. et al., "Facility Effects on

Stationary Plasma Thruster Testing," IEPC-93-

93, Sept. 1993.

12. Sankovic, J.M., Haag, T.W., and Manzella, D.H.,

"Operating Characteristics of the Russian D-55

Thruster with Anode Layer," AIAA-94-3011,

June 1994.

13. Sankovic, J.M., Haag, T.W., and Manzella, D.H.,

"Performance Evaluation of a 4.5 kW SPT

Thruster, " IEPC-95-30, Sept. 1995.

14. Kim, V. et al., "Development and Characterization of Small SPT," AIAA-98-3335, July

1998.

15. Manzella, D., Jacobson, D., and Jankovsky, R.,

"High Voltage SPT Performance," AIAA-2001-

3774, July 2001.

16. Hofer, R. and Jankovsky, R., "A Hall Thruster

Performance Model Incorporating the Effects of

a Multiply-Charged Plasma," AIAA-2001-3322,

July 2001.

17. Tilley, D., de Grys, K., and Myers, R., "Hall

Thruster--Cathode Coupling," AIAA-99-2865,

June 1999.

18. Jacobson, D., Jankovsky, R., and Manzella, D.,

"High Voltage TAL Performance," AIAA2001-3777.

NASA/TM--2002-211887 4
Figure 1: Photograph of the NASA-457M high power Hall thruster

1.4

1.2

1

0.8

0.6

0.4

0.2

0

-0.2

°_e_Boe,

L 1

--I nner - Calculated

Inner - Measured

-- Center - Calculated

-._-_ Center- Measured

•.......... Outer - Measured

Outer- Experiment

¢jAe_ °

a_ a

0.25 0.5 0.75 1 1.25 1.5

Normalized Axial distance from Anode

1.75 2

Figure 2a: Measured and predicted radial magnetic field strength

NASA/TM--2002-211887 5
0.6

0.4

0.2

0

-0.2

-0.4

-0.6

-0.8

-1

*°o

0 0.25 0.5 0.75 1 1.25

Normalized Axial distance from Anode

,_®O_O_Ae_OO_OQ

--Inner - Calculated

• Inner- Measured

- --Center - Calculated

,_ Center- Measured

............ Outer - Measured

_ Outer- Experiment

--[ F--

1.5 1.75

Figure 2b." Measured and predicted axial magnetic field strength

Figure3: Photograph test facility showing new test port.

NASA/TM--2002-211887 6
Figure 4: Photograph of thruster mounted on thrust stand in test port.

3.0

2.5

2.0

1.5

1.0

0.5

0.0

0

1

300 Volts

[W400 Volts

4500 Volts

_600 Volts

• 650 Volts

W

M

,&@ •

_o

¢

10 20 30 40 50 60 70

Discharge Power, kW

Figure 5: Thrust versus discharge power for various discharge voltages.

80

NASA/TM--2002-211887 7
3500

325O

3000

2750

2500

2250

2000

1750

• • • @

• • • o_ •

• <_-- -_ @ @

&

,d |

m _ mere t m [] Jl

R

300 Volts

[] 400 Volts --

,,&500 Volts

600 Volts __

• 650 Volts

1500

0 10 20 30 40 50 60 70

Discharge Power, kW

Figure 6: Discharge specific impulse versus discharge power for various discharge voltages.

8o

130

110 .--

90

I

300 Volts

[] 400 Volts

,_500 Volts

_600 Volts

• 650 Volts ,_

f

70

50

ly

30

10

0 10 20 30 40 50 60 70 80 90

Anode mass flow rate, mg/s

Figure 7: Discharge current versus anode mass flow rate for various discharge voltages.

100

NASA/TM--2002-211887 8
34OO

32OO

3OOO

_Model: Singly-Charged

_Model: Multiply-Charged

• Experimental Data

28OO

26OO

24OO

22OO

2OOO

f

300 400 500 600 700

Discharge Voltage, Volts

Figure 8: Predicted and measured specific impulse versus discharge voltage for an anode mass flow rate of 86.4 mg/s.

o

-30

-40

_ e _2.5 mg/s

-50 *_ E 5.0 mg/s '@

_7.5 mg/s

-60

-7O

10 20 30 40 50 60 70 80 90 100 110

Discharge Current, Amperes

Figure 9: Cathode-to-ground voltage versus discharge current for various cathode flow rates.

NASA/TM--2002-211887 9
0.70

0.65

0.60

0.55

0.50 [_

0.45

I1

e2.5 mg/s

_'5.0 mg/s

_7.5 mg/s

0.40

300 350 400 450 500 550 600 650

Discharge Voltage, V

FigurelO: Discharge efficiency versus discharge current for various cathode flow rates.

7OO

Anode Cathode

Discharge Discharge Discharge mass mass

voltage, current, power, flow, flow,

Volts Amperes Watts mg/s mg/s

300 34 10287 35.2 5.0

300 35 10383 35.2 2.5

300 36 10661 35.2 5.0

299 40 11880 40.4 5.0

298 41 12150 40.4 2.5

304 41 12480 40.4 2.5

301 41 12411 40.4 5.0

300 46 13740 45.8 5.0

302 47 14183 45.8 2.5

298 47 14012 45.8 2.5

301 47 14168 45.8 5.0

301 53 15974 51.3 5.0

300 53 16031 51.3 2.5

299 54 16010 51.3 5.0

300 58 17453 56.9 5.0

300 59 17622 56.9 5.0

299 60 18078 56.9 2.5

300 63 18971 62.6 7.5

300 65 19459 62.6 5.0

300 68 20290 62.6 5.0

299 71 21275 68.4 7.5

301 71 21481 68.4 7.5

300 72 21494 68.4 5.0

300 78 23422 74.3 7.5

300 79 23754 74.3 7.5

300 86 25894 80.3 7.5

301 88 26543 80.3 7.5

298 96 28587 86.4 7.5

300 99 29780 86.4 7.5

302 111 33454 92.7 7.5

Appendix: Data Table

Total Discharge

Total specific specific

power, Thrust, impulse, impulse, efficiency efficiency

Wa_s mN seconds seconds

11348 617 1567 1790

11483 614 1662 1781

11606 613 1557 1778

12961 701 1572 1766

13052 696 1653 1755

13735 721 1711 1817

13385 709 1590 1786

15010 812 1629 1807

15116 787 1661 1752

15015 798 1684 1776

15164 806 1617 1793

16999 904 1637 1796

16985 876 1660 1741

17358 930 1683 1847

18772 1005 1655 1801

18675 1005 1655 1801

19071 974 1671 1745

19643 1155 1679 1881

20542 1103 1663 1796

21664 1145 1726 1864

22400 1261 1694 1879

22829 1285 1726 1915

22596 1231 1710 1835

24701 1401 1746 1922

24901 1390 1732 1907

27194 1519 1763 1928

27707 1528 1774 1939

29923 1629 1767 1921

30970 1696 1840 2000

34663 1903 1937 2093

Total Discharge Cathode-to-ground

voltage,

Volts

0.42 0.53 -14.6

0.44 0.52 -37.9

0.40 0.50 -15.1

0.42 0.51 -17.3

0.43 0.49 -46.9

0.44 0.51 -45.6

0.41 0.50 -15.5

0.43 0.52 -17.9

0.42 0.48 -51.3

0.44 0.50 -45.4

0.42 0.50 -16.5

0.43 0.50 -19.9

0.42 0.47 -57.7

0.44 0.53 -18.9

0.43 0.51 -18.8

0.44 0.50 -21.5

0.42 0.46 -65.6

0.48 0.56 -11.4

0.44 0.50 -22.7

0.45 0.52 -22.8

0.47 0.55 -11.7

0.48 0.56 -11.6

0.46 0.52 -25

0.49 0.56 -12.4

0.47 0.55 -12.1

0.48 0.55 -12.8

0.48 0.55 -12.6

0.47 0.54 -12.9

0.49 0.56 -13.1

0.52 0.58 -13.7

NASA/TM--2002-211887 10
Appendix: Data Table (concluded)

Anode Cathode Total Discharge

Discharge Discharge Discharge mass

voltage, current, power, flow,

Volts Amperes Watts mg/s

403 25 9904 24.9

399 30 12016 30.0

400 35 14164 35.2

401 36 14399 35.2

403 37 14885 35.2

401 41 16561 40.4

404 43 17485 40.4

401 48 19349 45.8

397 50 19815 45.8

401 58 23080 51.3

400 61 24446 56.9

401 65 26159 62.6

399 74 29377 68.4

402 80 32144 74.3

402 89 35653 80.3

402 99 39697 86.4

501 20 9912 20.0

499 25 12517 24.9

499 31 15382 30.0

499 36 17950 35.2

500 37 18293 35.2

502 37 18570 35.2

501 42 21092 40.4

501 43 21497 40.4

500 44 22218 40.4

502 45 22649 40.4

500 50 25030 45.8

501 51 25601 45.8

504 59 29757 51.3

499 60 30064 51.3

500 65 32269 56.9

499 66 33154 62.6

497 76 37646 68.4

500 81 40484 74.3

499 90 44737 80.3

500 100 50120 86.4

500 100 50120 86.4

601 15 8956 15.1

599 20 12108 20.0

598 25 15075 24.9

599 30 17964 30.0

602 35 21056 35.2

600 40 24256 40.4

600 46 27637 45.8

600 52 31119 51.3

599 58 34431 56.9

599 64 38415 62.6

601 71 42777 68.4

600 79 47468 74.3

600 88 52680 80.3

602 97 58594 86.4

600 98 58500 86.4

599 108 64915 92.7

651 15 9706 15.1

650 20 13006 20.0

651 25 16022 24.9

652 29 19157 30.0

650 35 22562 35.2

650 41 26370 40.4

651 47 30811 45.8

651 53 34151 51.3

650 58 37878 56.9

650 65 42088 62.6

649 72 46915 68.4

649 80 51644 74.3

649 88 57242 80.3

649 99 63956 86.4

649 111 71963 92.7

mass Total

flow, power, Thrust, impulse, impulse, efficiency efficiency

specific specific Total Discharge Cathode-to-ground

voltage,

mg/s Wa_s mN seconds seconds Volts

2.5 10800 514 1911 2103 0.45 0.54 -24.4

2.5 13084 619 1942 2104 0.45 0.53 -31.7

5.0 15255 753 1912 2184 0.46 0.57 -14.4

2.5 15556 766 2072 2220 0.50 0.58 -40.5

2.5 16286 710 1921 2057 0.41 0.48 -48.3

5.0 17747 857 1922 2160 0.46 0.55 -17.4

2.5 18762 903 2143 2276 0.51 0.58 -46.8

5.0 20643 1010 2025 2247 0.49 0.58 -17.7

2.5 21055 992 2093 2207 0.48 0.54 -49.9

5.0 24442 1157 2095 2299 0.49 0.57 -19.5

5.0 25784 1258 2071 2253 0.50 0.57 -21

7.5 26867 1405 2044 2289 0.52 0.60 -11.4

7.5 30743 1538 2065 2292 0.51 0.59 -11.7

7.5 33446 1688 2103 2315 0.52 0.60 -12.4

7.5 36971 1815 2106 2303 0.51 0.57 -12.8

7.5 41050 2031 2204 2395 0.53 0.60 -13

2.5 10641 470 2131 2398 0.46 0.56 -19.2

2.5 13462 601 2233 2457 0.49 0.58 -24

2.5 16476 743 2330 2524 0.51 0.60 -31.7

5.0 19078 875 2220 2536 0.50 0.61 -14

5.0 19483 907 2301 2628 0.53 0.64 -13.8

2.5 19765 902 2442 2615 0.55 0.62 -42.4

5.0 22304 989 2220 2494 0.48 0.57 -16.8

5.0 22709 1062 2382 2676 0.55 0.65 -14.5

2.5 23323 1016 2412 2561 0.52 0.57 -42.2

2.5 23942 1062 2521 2677 0.55 0.62 -49

5.0 26341 1170 2346 2602 0.51 0.60 -17.5

2.5 26742 1166 2460 2595 0.53 0.58 -50.9

2.5 30932 1346 2550 2674 0.54 0.59 -53.1

5.0 31435 1358 2459 2698 0.52 0.60 -19.7

5.0 33631 1474 2428 2641 0.52 0.59 -21.2

7.5 33898 1617 2352 2633 0.55 0.63 -10.9

7.5 39028 1776 2385 2647 0.53 0.61 -11.7

7.5 41806 1930 2405 2648 0.54 0.62 -12.4

7.5 46074 2081 2415 2641 0.54 0.60 -12.8

7.5 51483 2330 2528 2747 0.56 0.63 -13.1

7.5 51483 2314 2511 2728 0.55 0.62 -13.1

5.0 10022 371 1881 2504 0.34 0.51 -11.6

5.0 13190 518 2113 2642 0.41 0.55 -13.6

5.0 16174 657 2237 2686 0.45 0.57 -14.4

5.0 19080 806 2347 2738 0.49 0.60 -14.6

5.0 22191 945 2400 2741 0.50 0.60 -15.4

5.0 25408 1090 2446 2748 0.51 0.61 -15.7

5.0 28801 1265 2538 2815 0.55 0.63 -16.5

5.0 32295 1398 2532 2778 0.54 0.61 -17.2

7.5 35626 1565 2477 2804 0.53 0.62 -10.2

7.5 39625 1735 2523 2826 0.54 0.63 -10.7

7.5 43996 1923 2583 2866 0.55 0.63 -11.2

7.5 48699 2121 2642 2909 0.56 0.64 -11.7

7.5 53923 2314 2685 2936 0.57 0.63 -12.2

7.5 59792 2524 2738 2976 0.57 0.63 -12.7

7.5 59755 2540 2756 2995 0.57 0.64 -12.7

10.0 66127 2778 2758 3055 0.57 0.64 -13.6

5.0 10599 398 2017 2684 0.37 0.54 -13.3

5.0 13920 541 2208 2761 0.42 0.56 -13.4

5.0 16953 688 2343 2813 0.47 0.59 -13.7

5.0 20115 834 2429 2834 0.49 0.60 -14.2

5.0 23541 969 2459 2809 0.50 0.59 -14.8

5.0 27368 1129 2534 2847 0.51 0.60 -15.2

5.0 31826 1288 2584 2866 0.51 0.59 -16.4

7.5 35231 1458 2527 2897 0.51 0.61 -9.7

7.5 38982 1621 2565 2904 0.52 0.61 -10.1

7.5 43213 1791 2605 2917 0.53 0.61 -10.6

7.5 48058 1983 2664 2956 0.54 0.61 -11.2

7.5 52840 2184 2722 2996 0.55 0.62 -11.6

7.5 58454 2409 2796 3057 0.57 0.63 -12.1

7.5 65182 2661 2888 3138 0.58 0.64 -12.8

10.0 73201 2950 2929 3245 0.58 0.65 -13.9

NASA/TM--2002-211887 11
Form Approved REPORT DOCUMENTATION PAGE

OMB No. 0704-0188

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

September 2002 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Laboratory Model 50 kW Hall Thruster

6. AUTHOR(S)

David Manzella, Robert Jankovsky, and Richard Hofer

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

DIJ-755-B4-05-00

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-13569

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM--2002-211887

AIAA-2002-3676

11. SUPPLEMENTARY NOTES

Prepared for the 38th Joint Propulsion Conference and Exhibit, cosponsored by the AIAA, ASME, SAE, and ASEE,

Indianapolis, Indiana, July 7-10, 2002. David Manzella, The University of Toledo 2801 W. Bancroft Street, Toledo,

Ohio 43606; Robert Jankovsky, NASA Glenn Research Center; and Richard Hofer QSS Group, Inc., Cleveland, Ohio

44135. Responsible person, David Manzella, organization code 5430, 216-977-7432.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category: 16 Distribution: Nonstandard

Available electronically at http://gltrs.grc.nasa.gov

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

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

A 0.46 meter diameter Hall thruster was fabricated and performance tested at powers up to 72 kilowatts. Thrusts

up to 2.9 Newtons were measured. Discharge specific impulses ranged from 1750 to 3250 seconds with discharge

efficiencies between 46 and 65 percent. Overall specific impulses ranged from 1550 to 3050 seconds with overall

efficiencies between 40 and 57 percent. Performance data indicated significant fraction of multiple-charged ions during

operation at elevated power levels. Cathode mass flow rate was shown to be a significant parameter with regard to

thruster efficiency.

14. SUBJECT TERMS

Electric propulsion; Hall thruster

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

NSN 7540-01-280-5500

15. NUMBER OF PAGES

17

16. PRICE CODE

18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT

OF THIS PAGE OF ABSTRACT

Unclassified Unclassified

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18

298-102