David Manzella

University of Toledo, Toledo, Ohio

Robert Jankovsky

Glenn Research Center, Cleveland, Ohio

Richard Hofer

QSS Group, Inc., Cleveland, Ohio

Laboratory Model 50 kW Hall Thruster

NASA/TM—2002-211887

September 2002

AIAA–2002–3676
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David Manzella

University of Toledo, Toledo, Ohio

Robert Jankovsky

Glenn Research Center, Cleveland, Ohio

Richard Hofer

QSS Group, Inc., Cleveland, Ohio

Laboratory Model 50 kW Hall Thruster

NASA/TM—2002-211887

September 2002

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Space Administration

Glenn Research Center

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NASA/TM—2002-211887 1

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 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. 

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.7

 

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 2

thruster, which has a 457 mm outer diameter ceramic 

discharge chamber, injects xenon propellant into the 

rear of the channel through a series of holes located 

along the inner and outer walls of the annular 

metallic anode located at the rear of the channel. The 

cathode, which has been previously described,9

 was 

located on the thruster centerline. 

A magnetic field is established across the exit plane 

of the discharge chamber through the use of two 

concentric electromagnets. Each electromagnet can 

be independently powered. The electromagnets along 

with magnetic poles and screens form the magnetic 

circuit. The design of the magnetic circuit was 

optimized through the use of a commercial threedimensional magneto-static computer code, Magnet 6 

by Infolytica. The field line topography employed in 

the NASA 457M is qualitatively similar to the 

plasma lens design of the NASA 173M.10 Following 

fabrication of the thruster the results from the 

magnetic model were compared with magnetic 

intensities measured with a 3-axis Gauss meter. The 

measured and calculated radial and axial components 

of the magnetic field are shown in figure 2. The 

channel depth normalized the axial distance and both 

the radial and axial magnetic field components were 

normalized by the maximum values respectively. 

The inner radial magnetic field strength was slightly 

higher than calculated which also affected the 

centerline radial field strength. This difference was 

attributed to uncertainty in the properties of the 

material used in this part of the magnetic circuit. 

The thruster was tested in a cryogenically pumped 

cylindrical vacuum chamber measuring 5 meters in 

diameter by 20 meters in length. This facility, which 

has been described in detail previously,10 was 

modified for this test. A cylindrical test port 

measuring 2 meters in diameter by 2.5 meters in 

length was mounted along the major axis of the tank 

with a 2-meter isolation valve between the test port 

and the main volume of the tank. A photograph of the 

test facility is shown in figure 3. The effective 

pumping speed of the test facility in this 

configuration was 700,000 liters per second. This was 

determined based on the xenon mass flow rate and 

the tank pressure measured behind the thruster. As a 

result, the highest background pressure was 2 × 10-5

Torr at a xenon flow rate of just over 1 standard liter 

per minute (SLPM). This pressure is still below the 

maximum recommended value for performance 

testing of 5 × 10-5 Torr.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 3

shown in figure 6 where specific impulse is plotted 

versus discharge power. Discharge specific impulse 

ranged from approximately 1750 seconds to 

3250 seconds. The overall thruster specific impulse 

including the cathode mass flow rate and magnet 

power ranged from 1550 seconds to 3050 seconds. 

While overall specific impulse and efficiency are 

reported, it should be noted that no attempts were 

made to optimize the laboratory model cathode or 

magnetic circuit for efficient operation. The 

discharge specific impulse at a given voltage 

compares favorably with other state-of-the-art 

thrusters. For example at 300 Volts an SPT-100 has a 

discharge specific impulse of 1750 seconds14 and 

the comparably sized high voltage SPT-1 thruster 

provided between 2250 and 2500 seconds at 

500 Volts.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 in the Hall thruster discharge 

channel is not fully understood and is more 

complicated than running a hollow cathode with a 

simple external anode due to the presence of 

magnetic fields. 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.15,18

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 4

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,” AIAA–

99–0457, 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, 27th 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,” IEPC–

93–151, 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,” AIAA–

2001–3777. 
NASA/TM—2002-211887 5

Figure 1: Photograph of the NASA-457M high power Hall thruster

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

0 0.25 0.5 0.75 1 1.25 1.5 1.75 2

Normalized Axial distance from Anode

Inner - Calculated

Inner - Measured

Center - Calculated

Center - Measured

Outer - Measured

Outer - Experiment

Figure 2a: Measured and predicted radial magnetic field strength
NASA/TM—2002-211887 6

-1

-0.8

-0.6

-0.4

-0.2

0

0.2

0.4

0.6

0 0.25 0.5 0.75 1 1.25 1.5 1.75 2

Normalized Axial distance from Anode

Inner - Calculated

Inner - Measured

Center - Calculated

Center - Measured

Outer - Measured

Outer - Experiment

Figure 2b: Measured and predicted axial magnetic field strength

Figure3: Photograph test facility showing new test port.
NASA/TM—2002-211887 7

Figure 4: Photograph of thruster mounted on thrust stand in test port.

0.0

0.5

1.0

1.5

2.0

2.5

3.0

0 10 20 30 40 50 60 70 80

Discharge Power, kW

300 Volts

400 Volts

500 Volts

600 Volts

650 Volts

Figure 5: Thrust versus discharge power for various discharge voltages.
NASA/TM—2002-211887 8

1500

1750

2000

2250

2500

2750

3000

3250

3500

0 10 20 30 40 50 60 70 80

Discharge Power, kW

300 Volts

400 Volts

500 Volts

600 Volts

650 Volts

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

10

30

50

70

90

110

130

0 10 20 30 40 50 60 70 80 90 100

Anode mass flow rate, mg/s

300 Volts

400 Volts

500 Volts

600 Volts

650 Volts

Figure 7: Discharge current versus anode mass flow rate for various discharge voltages.
NASA/TM—2002-211887 9

2000

2200

2400

2600

2800

3000

3200

3400

300 400 500 600 700

Discharge Voltage, Volts

Model: Singly-Charged

Model: Multiply-Charged

Experimental Data

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

-70

-60

-50

-40

-30

-20

-10

0

10 20 30 40 50 60 70 80 90 100 110

Discharge Current, Amperes

2.5 mg/s

5.0 mg/s

7.5 mg/s

Figure 9: Cathode-to-ground voltage versus discharge current for various cathode flow rates.
NASA/TM—2002-211887 10

0.40

0.45

0.50

0.55

0.60

0.65

0.70

300 350 400 450 500 550 600 650 700

Discharge Voltage, V

2.5 mg/s

5.0 mg/s

7.5 mg/s

Figure10: Discharge efficiency versus discharge current for various cathode flow rates.

Appendix: Data Table

Discharge

voltage,

Discharge

current,

Discharge

power,

Anode

mass

flow,

Cathode

mass

flow,

Total

power, Thrust,

Total

specific

impulse,

Discharge

specific

impulse,

Total

efficiency

Discharge

efficiency

Cathode-to-ground

voltage,

Volts Amperes Watts mg/s mg/s Watts mN seconds seconds Volts

300 34 10287 35.2 5.0 11348 617 1567 1790 0.42 0.53 -14.6

300 35 10383 35.2 2.5 11483 614 1662 1781 0.44 0.52 -37.9

300 36 10661 35.2 5.0 11606 613 1557 1778 0.40 0.50 -15.1

299 40 11880 40.4 5.0 12961 701 1572 1766 0.42 0.51 -17.3

298 41 12150 40.4 2.5 13052 696 1653 1755 0.43 0.49 -46.9

304 41 12480 40.4 2.5 13735 721 1711 1817 0.44 0.51 -45.6

301 41 12411 40.4 5.0 13385 709 1590 1786 0.41 0.50 -15.5

300 46 13740 45.8 5.0 15010 812 1629 1807 0.43 0.52 -17.9

302 47 14183 45.8 2.5 15116 787 1661 1752 0.42 0.48 -51.3

298 47 14012 45.8 2.5 15015 798 1684 1776 0.44 0.50 -45.4

301 47 14168 45.8 5.0 15164 806 1617 1793 0.42 0.50 -16.5

301 53 15974 51.3 5.0 16999 904 1637 1796 0.43 0.50 -19.9

300 53 16031 51.3 2.5 16985 876 1660 1741 0.42 0.47 -57.7

299 54 16010 51.3 5.0 17358 930 1683 1847 0.44 0.53 -18.9

300 58 17453 56.9 5.0 18772 1005 1655 1801 0.43 0.51 -18.8

300 59 17622 56.9 5.0 18675 1005 1655 1801 0.44 0.50 -21.5

299 60 18078 56.9 2.5 19071 974 1671 1745 0.42 0.46 -65.6

300 63 18971 62.6 7.5 19643 1155 1679 1881 0.48 0.56 -11.4

300 65 19459 62.6 5.0 20542 1103 1663 1796 0.44 0.50 -22.7

300 68 20290 62.6 5.0 21664 1145 1726 1864 0.45 0.52 -22.8

299 71 21275 68.4 7.5 22400 1261 1694 1879 0.47 0.55 -11.7

301 71 21481 68.4 7.5 22829 1285 1726 1915 0.48 0.56 -11.6

300 72 21494 68.4 5.0 22596 1231 1710 1835 0.46 0.52 -25

300 78 23422 74.3 7.5 24701 1401 1746 1922 0.49 0.56 -12.4

300 79 23754 74.3 7.5 24901 1390 1732 1907 0.47 0.55 -12.1

300 86 25894 80.3 7.5 27194 1519 1763 1928 0.48 0.55 -12.8

301 88 26543 80.3 7.5 27707 1528 1774 1939 0.48 0.55 -12.6

298 96 28587 86.4 7.5 29923 1629 1767 1921 0.47 0.54 -12.9

300 99 29780 86.4 7.5 30970 1696 1840 2000 0.49 0.56 -13.1

302 111 33454 92.7 7.5 34663 1903 1937 2093 0.52 0.58 -13.7
NASA/TM—2002-211887 11

Appendix: Data Table (concluded)

Discharge

voltage,

Discharge

current,

Discharge

power,

Anode

mass

flow,

Cathode

mass

flow,

Total

power, Thrust,

Total

specific

impulse,

Discharge

specific

impulse,

Total

efficiency

Discharge

efficiency

Cathode-to-ground

voltage,

Volts Amperes Watts mg/s mg/s Watts mN seconds seconds Volts

403 25 9904 24.9 2.5 10800 514 1911 2103 0.45 0.54 -24.4

399 30 12016 30.0 2.5 13084 619 1942 2104 0.45 0.53 -31.7

400 35 14164 35.2 5.0 15255 753 1912 2184 0.46 0.57 -14.4

401 36 14399 35.2 2.5 15556 766 2072 2220 0.50 0.58 -40.5

403 37 14885 35.2 2.5 16286 710 1921 2057 0.41 0.48 -48.3

401 41 16561 40.4 5.0 17747 857 1922 2160 0.46 0.55 -17.4

404 43 17485 40.4 2.5 18762 903 2143 2276 0.51 0.58 -46.8

401 48 19349 45.8 5.0 20643 1010 2025 2247 0.49 0.58 -17.7

397 50 19815 45.8 2.5 21055 992 2093 2207 0.48 0.54 -49.9

401 58 23080 51.3 5.0 24442 1157 2095 2299 0.49 0.57 -19.5

400 61 24446 56.9 5.0 25784 1258 2071 2253 0.50 0.57 -21

401 65 26159 62.6 7.5 26867 1405 2044 2289 0.52 0.60 -11.4

399 74 29377 68.4 7.5 30743 1538 2065 2292 0.51 0.59 -11.7

402 80 32144 74.3 7.5 33446 1688 2103 2315 0.52 0.60 -12.4

402 89 35653 80.3 7.5 36971 1815 2106 2303 0.51 0.57 -12.8

402 99 39697 86.4 7.5 41050 2031 2204 2395 0.53 0.60 -13

501 20 9912 20.0 2.5 10641 470 2131 2398 0.46 0.56 -19.2

499 25 12517 24.9 2.5 13462 601 2233 2457 0.49 0.58 -24

499 31 15382 30.0 2.5 16476 743 2330 2524 0.51 0.60 -31.7

499 36 17950 35.2 5.0 19078 875 2220 2536 0.50 0.61 -14

500 37 18293 35.2 5.0 19483 907 2301 2628 0.53 0.64 -13.8

502 37 18570 35.2 2.5 19765 902 2442 2615 0.55 0.62 -42.4

501 42 21092 40.4 5.0 22304 989 2220 2494 0.48 0.57 -16.8

501 43 21497 40.4 5.0 22709 1062 2382 2676 0.55 0.65 -14.5

500 44 22218 40.4 2.5 23323 1016 2412 2561 0.52 0.57 -42.2

502 45 22649 40.4 2.5 23942 1062 2521 2677 0.55 0.62 -49

500 50 25030 45.8 5.0 26341 1170 2346 2602 0.51 0.60 -17.5

501 51 25601 45.8 2.5 26742 1166 2460 2595 0.53 0.58 -50.9

504 59 29757 51.3 2.5 30932 1346 2550 2674 0.54 0.59 -53.1

499 60 30064 51.3 5.0 31435 1358 2459 2698 0.52 0.60 -19.7

500 65 32269 56.9 5.0 33631 1474 2428 2641 0.52 0.59 -21.2

499 66 33154 62.6 7.5 33898 1617 2352 2633 0.55 0.63 -10.9

497 76 37646 68.4 7.5 39028 1776 2385 2647 0.53 0.61 -11.7

500 81 40484 74.3 7.5 41806 1930 2405 2648 0.54 0.62 -12.4

499 90 44737 80.3 7.5 46074 2081 2415 2641 0.54 0.60 -12.8

500 100 50120 86.4 7.5 51483 2330 2528 2747 0.56 0.63 -13.1

500 100 50120 86.4 7.5 51483 2314 2511 2728 0.55 0.62 -13.1

601 15 8956 15.1 5.0 10022 371 1881 2504 0.34 0.51 -11.6

599 20 12108 20.0 5.0 13190 518 2113 2642 0.41 0.55 -13.6

598 25 15075 24.9 5.0 16174 657 2237 2686 0.45 0.57 -14.4

599 30 17964 30.0 5.0 19080 806 2347 2738 0.49 0.60 -14.6

602 35 21056 35.2 5.0 22191 945 2400 2741 0.50 0.60 -15.4

600 40 24256 40.4 5.0 25408 1090 2446 2748 0.51 0.61 -15.7

600 46 27637 45.8 5.0 28801 1265 2538 2815 0.55 0.63 -16.5

600 52 31119 51.3 5.0 32295 1398 2532 2778 0.54 0.61 -17.2

599 58 34431 56.9 7.5 35626 1565 2477 2804 0.53 0.62 -10.2

599 64 38415 62.6 7.5 39625 1735 2523 2826 0.54 0.63 -10.7

601 71 42777 68.4 7.5 43996 1923 2583 2866 0.55 0.63 -11.2

600 79 47468 74.3 7.5 48699 2121 2642 2909 0.56 0.64 -11.7

600 88 52680 80.3 7.5 53923 2314 2685 2936 0.57 0.63 -12.2

602 97 58594 86.4 7.5 59792 2524 2738 2976 0.57 0.63 -12.7

600 98 58500 86.4 7.5 59755 2540 2756 2995 0.57 0.64 -12.7

599 108 64915 92.7 10.0 66127 2778 2758 3055 0.57 0.64 -13.6

651 15 9706 15.1 5.0 10599 398 2017 2684 0.37 0.54 -13.3

650 20 13006 20.0 5.0 13920 541 2208 2761 0.42 0.56 -13.4

651 25 16022 24.9 5.0 16953 688 2343 2813 0.47 0.59 -13.7

652 29 19157 30.0 5.0 20115 834 2429 2834 0.49 0.60 -14.2

650 35 22562 35.2 5.0 23541 969 2459 2809 0.50 0.59 -14.8

650 41 26370 40.4 5.0 27368 1129 2534 2847 0.51 0.60 -15.2

651 47 30811 45.8 5.0 31826 1288 2584 2866 0.51 0.59 -16.4

651 53 34151 51.3 7.5 35231 1458 2527 2897 0.51 0.61 -9.7

650 58 37878 56.9 7.5 38982 1621 2565 2904 0.52 0.61 -10.1

650 65 42088 62.6 7.5 43213 1791 2605 2917 0.53 0.61 -10.6

649 72 46915 68.4 7.5 48058 1983 2664 2956 0.54 0.61 -11.2

649 80 51644 74.3 7.5 52840 2184 2722 2996 0.55 0.62 -11.6

649 88 57242 80.3 7.5 58454 2409 2796 3057 0.57 0.63 -12.1

649 99 63956 86.4 7.5 65182 2661 2888 3138 0.58 0.64 -12.8

649 111 71963 92.7 10.0 73201 2950 2929 3245 0.58 0.65 -13.9
This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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Technical Memorandum

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National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135–3191

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September 2002

NASA TM—2002-211887

AIAA–2002–3676

E–13569

WU–755–B4–05–00

17

Laboratory Model 50 kW Hall Thruster

David Manzella, Robert Jankovsky, and Richard Hofer

Electric propulsion; Hall thruster

Unclassified -Unlimited

Subject Category: 16 Distribution: Nonstandard

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.

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.