George C. Soulas, Thomas W. Haag, Daniel A. Herman, Wensheng Huang,

Hani Kamhawi, and Rohit Shastry

Glenn Research Center, Cleveland, Ohio

Performance Test Results of the NASA-457M v2 

Hall Thruster

NASA/TM—2012-217711

August 2012

AIAA–2012–3940
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George C. Soulas, Thomas W. Haag, Daniel A. Herman, Wensheng Huang,

Hani Kamhawi, and Rohit Shastry

Glenn Research Center, Cleveland, Ohio

Performance Test Results of the NASA-457M v2 

Hall Thruster

NASA/TM—2012-217711

August 2012

AIAA–2012–3940

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

48th Joint Propulsion Conference and Exhibit

cosponsored by AIAA, ASME, SAE, and ASEE

Atlanta, Georgia, July 30 to August 1, 2012
Acknowledgments

The authors acknowledge David M. Manzella, David T. Jacobson, and Robert S. Jankovsky for the design and fabrication of the 

NASA-457M v2 Hall thruster. The authors also acknowledge Kevin L. Blake, George W. Readus, and George P. Jacynycz for the 

fabrication of the thruster, assembly of the test setup, and operation of the vacuum facility.

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Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2012-217711 1 

Performance Test Results of the NASA-457M v2 Hall Thruster 

George C. Soulas, Thomas W. Haag, Daniel A. Herman, Wensheng Huang, 

Hani Kamhawi, and Rohit Shastry 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

Performance testing of a second generation, 50 kW-class Hall thruster labeled NASA-457M v2 was 

conducted at the NASA Glenn Research Center. This NASA-designed thruster is an excellent candidate 

for a solar electric propulsion system that supports human exploration missions. Thruster discharge power 

was varied from 5 to 50 kW over discharge voltage and current ranges of 200 to 500 V and 15 to 100 A, 

respectively. Anode efficiencies varied from 0.56 to 0.71. The peak efficiency was similar to that of other 

state-of-the-art high power Hall thrusters, but outperformed these thrusters at lower discharge voltages. 

The 0.05 to 0.18 higher anode efficiencies of this thruster compared to its predecessor were primarily due 

to which of two stable discharge modes the thruster was operated. One stable mode was at low magnetic 

field strengths, which produced high anode efficiencies, and the other at high magnetic fields where its 

predecessor was operated. Cathode keeper voltages were always within 2.1 to 6.2 V and cathode voltages 

were within 13 V of tank ground during high anode efficiency operation. However, during operation at 

high magnetic fields, cathode-to-ground voltage magnitudes increased dramatically, exceeding 30 V, due 

to the high axial magnetic field strengths in the immediate vicinity of the centrally-mounted cathode. The 

peak thrust was 2.3 N and this occurred at a total thruster input power of 50.0 kW at a 500 V discharge 

voltage. The thruster demonstrated a thrust-to-power range of 76.4 mN/kW at low power to 46.1 mN/kW 

at full power, and a specific impulse range of 1420 to 2740 s. For a discharge voltage of 300 V, where 

specific impulses would be about 2000 s, thrust efficiencies varied from 0.57 to 0.63. 

Nomenclature 

GRC Glenn Research Center 

HEFT Human Exploration Framework Team 

ISP In Space Propulsion Program 

J current 

Jd discharge current 

݉ሶ

௔ anode mass flow 

T thrust 

V voltage 

Vd discharge voltage 

ηa anode efficiency 

Introduction 

In 2010, NASA established the Human Exploration Framework Team (HEFT) to analyze exploration 

and technology concepts and provide inputs to the agency’s senior leadership on the key components of a 

safe, sustainable, affordable, and credible future human space exploration endeavor for the nation 

(Ref. 1). The team concluded, in part, that the use of a high power (i.e., on the order of 300 kW) solar 

electric propulsion system could significantly reduce the number of heavy lift launch vehicles required for 
NASA/TM—2012-217711 2 

a human mission to a near earth asteroid (Refs. 1 and 2). Hall thrusters were found to be ideal for such 

applications because of their high power processing capabilities and their efficient operation at moderate 

specific impulses, which leads to reduced trip times that are necessary for such missions (Refs. 2 to 4). 

Recent electric propulsion system model estimates that considered factors such as cost, mass, faulttolerance, cost uncertainty, complexity, ground test vacuum facility limitations, previously demonstrated 

power capabilities, and possible technology limitations have shown that Hall thrusters operating at power 

levels of 20 to 50 kW are strong candidates for human exploration missions operating at total powers up 

to 500 kW (Refs. 2 and 5). 

An excellent thruster candidate for such a mission is the NASA-457M Hall thruster. This thruster is a 

50 kW-class thruster whose development was initiated in 2000 for the NASA Space Solar Power Concept 

and Technology Maturation Program to enable space solar power systems and other high power 

spacecraft (Refs. 6 and 7). A laboratory model version of this high power thruster was fabricated and 

tested with xenon in 2002 at power levels up to 72 kW and at discharge voltages between 300 and 

650 V (Refs. 7 and 8). Anode specific impulses were measured to be between 1750 to 3250 s with anode 

efficiencies between 0.46 and 0.65 on xenon (Refs. 7 and 9). Later tests with this thruster were conducted 

with a krypton propellant over a similar input power range but with discharge voltages up to 1000 V to 

achieve even higher specific impulses (Ref. 10). This thruster was further tested with xenon at low 

discharge voltages (100 to 150 V) to evaluate performance at the low specific impulse range of 800 to 

1500 s for near-earth space applications requiring high thrust-to-power (Refs. 8 and 11). 

The successful test campaign with the laboratory model NASA-457M led to the development of a 

higher fidelity version of this thruster, labeled the NASA-457M v2. This work was funded by NASA’s 

In-Space Transportation Program, whose objective was to advance the technology readiness level of highpower Hall thruster technology (Ref. 12). Thruster fabrication was completed in 2004 and initial checkout 

testing was initiated thereafter. Unfortunately, program funding was terminated prior to the initiation of 

performance tests. 

Testing of the NASA-457M v2 was re-initiated last fiscal year with funding from NASA’s Human 

Exploration and Operations Mission Directorate Enabling Technology Development and Demonstration 

Program (Ref. 13). This program was focused on developing and maturing the technologies needed to 

reduce cost and expand the capability of future space exploration activities. This fiscal year, this program 

was transitioned to the NASA’s Office of Chief Technologist Game Changing Technology Division and 

is called the In Space Propulsion (ISP) project. The program content is still focused on developing and 

maturing the high power propulsion technologies needed to enhance the agency’s capabilities to explore 

and move large payloads in space (Ref. 13). Testing of this thruster was re-initiated with funding from 

both programs. 

The test program for the NASA-457M v2 had multiple objectives. These objectives are listed below 

along with their impacts: 

1. To evaluate the performance of the NASA-457M v2 over the full input power range of 5 to 50 kW 

with a xenon propellant. This thruster could provide the ISP program with a potential head start in 

high power Hall thruster development if the thruster meets future mission requirements. 

2. Along with NASA-300M thruster performance testing (Ref. 13), to reconstitute a high power Hall 

thruster testing capability at the NASA Glenn Research Center (GRC). A successful test completion 

would provide NASA’s ISP program with a test stand for future high power thruster testing. 

3. To provide test data for: a) verification of thruster thermal and service life models, as well as vacuum 

facility models; and b) development of in situ erosion diagnostics. Successful model verification and 

erosion diagnostics development would decrease thruster development costs and risks for the ISP 

program by reducing the number of design iterations and service life tests. 

4. To provide high power thruster operating characteristics for power processor development. These 

data would help reduce future power processor development risks. 
NASA/TM—2012-217711 3 

This paper will only present the performance test results of the NASA-457M v2 and compare them 

to those of the laboratory model. The thruster and test facility will initially be described. The section 

thereafter will present the performance test results and compare them to the laboratory model 

NASA-457M.

Apparatus and Procedures 

Thruster 

A photograph of the higher fidelity 50 kW Hall thruster (i.e., the NASA-457M v2), is shown in 

Figure 1. This high power thruster has a solid boron nitride discharge chamber with a 457 mm outer 

diameter and a centrally-mounted high current hollow cathode. The development goal for the 

NASA-457M v2 was to advance its technology readiness level to 4-5. The fidelity of the laboratory 

model NASA-457M was improved by addressing thermal and mechanical issues while maintaining the 

critical features of the laboratory model thruster (Ref. 12). The higher fidelity thruster retained the same 

discharge chamber critical dimensions, however the magnetic circuit was designed for improved 

performance and field symmetry. The field topology employed in the NASA 457M v2 is very similar to 

that of the laboratory model field, which incorporated the plasma lens topography (Refs. 14 and 15). The 

centerline radial magnetic field strength was increased by nearly 20 percent for the same number of Ampturns (Ref. 12). The magnetic fields produced by the NASA-457M v2 thruster were measured and 

compared to its laboratory model predecessor. Although measured magnetic fields for the laboratory 

model thruster were already reported in References 7 and 8, they were re-measured for these 

comparisons. The measured field strengths of both thrusters were found to be similar. 

In addition to the magnetic field circuit, the new mechanical design eliminated deficiencies with 

respect to anode mounting and electrical isolation, concentricity, and thermally-induced mechanical 

interferences (Ref. 12). The thruster’s design was assessed using finite element thermal and mechanical 

modeling techniques. Mechanical analyses assessed vibration tolerance, static deflection, and induced 

stresses. Thermal models were applied to the mechanical models to predict thermal displacements and 

corresponding stresses. The thruster’s design was updated based on the results of these analyses. 

Although a high emission current hollow cathode was also under development (Ref. 12), a laboratory 

model hollow cathode assembly was used for all performance testing reported herein. This cathode 

assembly is the same design as that used for performance tests of the laboratory model thruster (Refs. 7, 

10, and 11). The cathode assembly design is based on the discharge cathode assembly design used for 

NASA’s Evolutionary Xenon Thruster with adjustments to cathode and keeper orifice diameters to 

account for the higher emission currents (Ref. 16). 

Figure 1.—Photograph of the 

NASA-457M v2 thruster. 
NASA/TM—2012-217711 4 

Vacuum Facility and Test Support Hardware 

As was done with the laboratory model version thruster, performance testing was conducted in NASA 

GRC’s Vacuum Facility 5. This cylindrical facility is 4.6 m in diameter and 18.3 m long and was pumped 

with cryogenically-cooled panels and 20 axially distributed oil diffusion pumps. Unlike the laboratory 

model thruster test (Ref. 7), the thruster was mounted in the main volume of the vacuum facility to ensure 

low pressures near the thruster. A photograph of the mounted thruster is shown in Figure 2. Facility 

pressures were measured with an ion gage mounted next to the thruster’s thrust stand mount. Facility 

pressure uncertainties were estimated by the manufacturer to be within 6 percent of reading. Pressures 

next to the thruster were less than 4.0×10–3 Pa (3.0×10–5 torr) throughout performance testing. 

A gas feed system made of commercially available components was used to feed xenon to the 

thruster. The gas feed system utilized 1000 and 200 sccm mass flow controllers to provide propellant to 

the anode and cathode, respectively. Both mass flow controllers were calibrated prior to testing using 

xenon. Typical uncertainties reported herein were 1.0 percent of reading for the anode mass flow and 

1.1 percent of reading for total mass flows. 

A commercial discharge power supply was used and was capable of supplying 2000 V at over 100 A 

with a 5.5 mF output filter capacitance. A 15.3 mF capacitor bank was connected in parallel with the 

power supply at the electrical feed-throughs of the vacuum facility to filter discharge current oscillations. 

In addition, separate commercial power supplies were used to power the cathode heater, cathode keeper, 

and electromagnets. 

A null-type inverted pendulum thrust stand was used during performance testing, which is identical to 

the one used in References 7 and 8. The power cables were fed to the thruster using a “water fall” 

configuration to minimize thermal drift (Ref. 16). In-situ thrust stand calibrations were performed prior to 

and after all performance tests. In addition, the thruster was periodically powered off during performance 

testing to measure the thermal drift, which was used to correct the thrust measurements reported herein. 

Thrust uncertainty was estimated to be 2 percent of reading (Refs. 7 and 16). 

Figure 2.—Photograph of the NASA-457M 

v2 thruster mounted in NASA’s Vacuum 

Facility 5. 

 
NASA/TM—2012-217711 5 

A data logger was used to measure and record thruster currents, voltages, and mass flow rates, as well as 

facility pressures. All data logger measurements were calibrated prior to testing using calibrated meters. 

Current uncertainties were estimated to be within 0.3 percent of reading for discharge currents and within 

0.6 percent of reading for all other currents. Voltage uncertainties were estimated to be within 0.1 percent 

of reading for discharge voltages and within 0.04 percent of reading for all other voltages. 

The test also included a number of other diagnostics to address the other test objectives listed earlier. 

These diagnostics are listed below: 

1. Optical diagnostics were included to support the development of an erosion diagnostic. The thruster 

downstream insulator was interrogated for neutral insulator material, as well as neutral, singlycharged, and doubly-charged xenon with seven stationary lenses coupled to two spectrometers via 

fiber optic cables. The near-field plume was also interrogated for neutral and singly-charged xenon 

with separate cameras. 

2. Near- and far-field plasma plume diagnostics were installed on a two axis positioning system to 

support thruster model development. They included a planar probe for beam current densities, a 

single Langmuir probe for electron temperatures and number densities, emissive probes for plasma 

potentials, and Wien filters (more commonly called ExB probes) for charge states. 

3. Thermocouples were mounted on the outer ceramic channel wall and on the inner and outer 

ferromagnetic poles to support thermal model development. 

4. Discharge current and voltage probes were used to measure oscillations to support power processor 

development. 

5. Four additional vacuum facility ion gages were mounted at various locations to support facility 

pressure model development. 

Results from these diagnostics will not be reported herein, but most will be the topic of future 

publications. 

Throttling Range and Operating Procedures 

The discharge currents, voltages, and powers of this performance test are listed in Table 1. Thruster 

discharge power was varied from 5 to 50 kW over a 200 to 500 V discharge voltage range. The maximum 

discharge voltage was set by the maximum anticipated specific impulse required by the HEFT mission, 

which favored values less than about 2600 s to keep mission durations acceptable (Ref. 2). A discharge 

voltage of 200 V was selected as the minimum because of the significant decrease in anode efficiencies as 

a function of discharge voltage. Although performance was originally intended to be measured in 100 V 

increments, performance was further measured at 250 V because of the large change in anode efficiency 

between the 200 and 300 V cases. Performance was also measured at 350 V to support erosion 

diagnostics development. 

The maximum discharge current was limited to 100 A because the cathode was not designed to 

operate at such high currents for extended periods and the discharge power supply was nearing its high 

current capability. A minimum discharge current of 15 A was selected because of the significant decrease 

of anode efficiencies as a function of discharge current at that level. 

Before the start of performance testing, thruster input power was increased gradually as a function of 

time to slowly remove moisture absorbed by the discharge channel’s boron nitride, which ultimately 

permitted stable thruster operation at high powers. This process is referred to as insulator bakeout. 

Following insulator bakeout, performance was measured over a several day period. For the first throttle 

level of a given day, the thruster was operated for at least 1 h before recording thruster performance. 

Subsequent throttle levels were operated until performance had achieved equilibrium, which typically 

required at least 10 min of operation at that level. Throughout performance testing, the cathode keeper 

current was set to 1.0 A and the inner and outer magnet currents were always set equal to each other. 
NASA/TM—2012-217711 6 

TABLE 1.—DISCHARGE THROTTLE LEVELS FOR THE 

NASA-457M V2 THRUSTER PERFORMANCE TEST 

Discharge 

power, 

kW

Discharge current, A vs. Discharge voltage, V

200 V 250 V 300 V 350 V 400 V 500 V 

5 a25 20 16.7 

6 - - - - 15 

7.5 - - - - - 15 

10 a50 40 33.3 - a25 - 

15 a75 60 50 - 37.5 30 

20 a100 80 66.7 57.1 a50 40 

25 100 83.3 - 62.5 50 

30 100 85.7 a75 60 

35 - 87.5 70 

40 a100 80 

45 90 

50 100 a

Indicates high magnetic field operation was conducted in addition to high anode 

efficiency operation.

The thruster was operated in two stable modes: high anode efficiency and high magnetic field 

operation. Stable thruster operation is defined here as peak-to-peak discharge current oscillations less than 

that of the average discharge current. During high anode efficiency operation, anode flows and magnetic 

fields were adjusted to optimize anode efficiency for all discharge voltages and currents listed in Table 1. 

Here, anode efficiency, ηa, is defined as: 

dda a VJm

T

  2 

2

 (1) 

where T is the measured thrust, ma is the anode flow, and Jd and Vd are the discharge current and voltage, 

respectively. During anode efficiency optimization, peak-to-peak discharge current oscillations were 

typically  80 percent of the average discharge current. The cathode flow was also set to 8.0 percent of 

the anode flow. Most of the performance data presented herein were measured with the thruster operated 

in this manner. 

During high magnetic field operation, the thruster was operated at a constant maximum radial 

magnetic field along the discharge channel centerline for the limited throttling points indicated in Table 1. 

This radial magnetic field setting was higher than the highest magnetic field setting during high anode 

efficiency operation. This operating mode simulated the manner in which the NASA-457M laboratory 

model thruster was operated in References 7 and 8 so that comparisons between the two thruster versions 

could be made. For this operating mode, the cathode flow was set to values similar to those of 

References 7 and 8, and 8.0 percent of the anode flow. 

Results and Discussions 

The following sections present and discuss the results of the performance test. The initial sections 

present and discuss the discharge and cathode performance results separately. Overall thruster performance 

is presented and discussed thereafter. In all sections, the results are interpreted and compared to that of the 

laboratory model version of the thruster, whose performance was reported in References 7 and 8. 
NASA/TM—2012-217711 7 

Discharge Performance 

Anode efficiency as a function of discharge power for the v2 thruster is plotted in Figure 3 for high 

anode efficiency operation. The uncertainty in discharge power for all results reported herein was 

estimated to be 0.3 percent of the indicated value and anode efficiency was estimated to typically be 

4.2 percent of the indicated value. As the figure shows, anode efficiencies varied from 0.56 to 0.71 and 

typically increased with increasing discharge voltage. The peak efficiency achieved by the NASA-457M 

v2 was similar to that of other high power Hall thrusters, such as the NASA-400M, the NASA-300M, and 

the BHT-20K which achieved peak efficiencies of 0.71, 0.73, and 0.68, respectively (Refs. 16 to 18). The 

NASA-457M v2, however, outperformed these other thrusters at lower discharge voltages. For example, 

the NASA-457M v2 demonstrated an anode efficiency range of 0.56 to 0.63 at a discharge voltage of 

200 V, while the NASA-400M, the NASA-300M, and the BHT-20K demonstrated anode efficiency 

ranges of 0.46 to 0.49, 0.51 to 0.57, and 0.51 to 0.54, respectively (Refs. 16 to 18). 

Figure 3 also shows that anode efficiency as a function of discharge power generally increased with 

power at lower powers at each discharge voltage, achieved a peak value at an intermediate power level, 

and then decreased with power at higher powers. For comparison, anode efficiency as a function of 

discharge power at various discharge voltages for the laboratory model thruster from References 7 and 8 

is plotted in Figure 4. A comparison of the two figures shows two distinct differences. First, the higher 

fidelity thruster appears to outperform the laboratory model thruster by a difference of 0.05 to 0.18. 

Second, the laboratory model anode efficiency as a function of discharge power either increased with 

discharge power, or increased with power at lower powers and then achieved a plateau value for 

discharge voltages greater than 200 V. 

The main reason for the anode efficiency differences between the two thruster versions was 

differences in how they were operated. The laboratory model thruster was operated with a nearly constant 

peak centerline radial magnetic field. During performance testing of the v2 thruster, magnetic fields were 

adjusted to optimize anode efficiency. During this optimization, two stable modes of operation were 

found. One was at low magnetic field strengths, which produced high anode efficiencies, and the other at 

high magnetic fields. 

Figure 3.—Anode efficiency as a function of discharge power at various discharge 

voltages for the NASA-457M v2 during operation in high anode efficiency mode. 

0.35

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0 10,000 20,000 30,000 40,000 50,000

Anode Efficiency

Discharge Power, W

500 V

400 V

300 V

250 V

200 V
NASA/TM—2012-217711 8 

Figure 4.—Anode efficiency as a function of discharge power at various 

discharge voltages for the laboratory model NASA-457M. Data are from 

References 7 and 8. 

Figure 5.—Anode efficiency as a function of discharge power for the laboratory 

model and v2 thrusters at high magnetic fields and high anode efficiencies. 

Laboratory model data are from References 7 and 8. 

To permit a more direct comparison between the two thruster versions, the v2 thruster was operated at 

a constant peak radial magnetic field along the discharge channel centerline that was about 75 percent that 

of the laboratory model thruster. Anode efficiencies as a function of discharge power are shown in 

Figure 5 for both thrusters at two discharge voltages. The figure shows that when the v2 thruster’s 

magnetic fields were increased, anode efficiencies decreased between 0.05 and 0.11, with the 

performance of the v2 thruster approaching that of the laboratory model thruster. Indeed, the performance 

differences between the two thrusters at high magnetic fields were within measurement uncertainties. In 

addition, anode efficiency decreases at higher powers were less pronounced. It can be concluded from 

Figures 3 to 5 that anode performance between the two thruster versions was quite similar, and that the 

NASA-457M could be operated at lower magnetic field strengths to achieve more optimal anode 

efficiencies than had been done in the past. 

0.35

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0 10,000 20,000 30,000 40,000 50,000

Anode Efficiency

Discharge Power, W

500 V

400 V

300 V

200 V

0.35

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0 10,000 20,000 30,000 40,000 50,000

Anode Efficiency

Discharge Power, W

400 V ‐ v2 High Anode Efficiency

400 V ‐ v2 High Magnetic Field

400 V ‐ Lab Model

200 V ‐ v2 High Anode Efficiency

200 V ‐ v2 High Magnetic Field

200 V ‐ Lab Model
NASA/TM—2012-217711 9 

The reason for the lower anode efficiencies at high magnetic fields is explored in Figures 6 and 7, 

which plot thrust as a function of discharge current and discharge current as a function of anode mass 

flow, respectively, for the two thrusters operating in different modes. As shown in Equation (1), anode 

efficiency is directly proportional to thrust squared and indirectly proportional to anode mass flow and 

discharge current. Figures 6 and 7 demonstrate that anode efficiencies were lower at high magnetic fields 

because, in general, thrusts were lower. However, Figure 6 also shows that this thrust disparity decreased 

as a function of increasing discharge current for currents  60 A, until thrusts were equal at about 100 A. 

Regardless, operation with lower magnetic fields still produced higher anode efficiencies for discharge 

currents  60 A because the thruster was able achieve those thrusts with lower anode mass flows, as 

shown in Figure 7. 

Figure 6.—Thrust as a function of discharge current for the v2 thruster during 

high anode efficiency operation and the laboratory model thruster during 

operation with high magnetic fields at discharge voltages of 200 and 

400 V. Laboratory model data are from References 7 and 8. 

Figure 7.—Discharge current as a function of anode mass flow for the v2 

thruster during high anode efficiency operation and the laboratory model 

thruster during operation with high magnetic fields. Laboratory model data 

are from References 7 and 8. 

0

500

1000

1500

2000

2500

10 20 30 40 50 60 70 80 90 100 110

Thrust, mN

Discharge Current, A

400 V ‐ v2;High Anode Eff.

400 V ‐ Lab Model

200 V ‐ v2;High Anode Eff.

200 V ‐ Lab Model

10

20

30

40

50

60

70

80

90

100

110

120

10 20 30 40 50 60 70 80 90 100

DischargeCurrent, A

AnodeMass Flow, mg/s

v2;High Anode Eff.

Lab Model
NASA/TM—2012-217711 10 

This was particularly unexpected because minimizing the discharge current for a given anode flow, 

and therefore a lower electron current to the anode, is typically expected to yield a more optimal anode 

efficiency (Ref. 19). The nonlinear behavior of v2 thruster thrust as a function of discharge current for 

currents  60 A in Figure 6, and its increasing discharge current in Figure 7 as a function of anode mass 

flow, suggest that beam ion current is decreasing for currents  60 A during high anode efficiency 

operation. And the integrated beam current densities for the NASA-457M v2 thruster operating at a 

discharge current and voltage of 100 A and 200 V, respectively, showed that beam ion currents were 

indeed higher during high magnetic field operation (Ref. 20). However, even though the laboratory model 

thruster operating with high magnetic fields likely produced a higher ion beam current for discharge 

currents  60 A as shown in Figure 7, the thrust was still less than or equal to that of the v2 thruster at 

lower magnetic fields as shown in Figure 6. This suggests that other thrust loss mechanisms are becoming 

significant during operation with high magnetic fields, such as beam divergence, the ratio of singly-tomultiply charged ions, or coupling voltage losses. In summary, the reason for the lower anode efficiencies 

at high magnetic fields was lower thrusts, even when data suggested that higher magnetic fields were 

producing higher beam ion currents. 

The reason for the anode efficiency trends with discharge power can be better understood by plotting 

anode efficiency as a function of discharge current, shown in Figure 8 for the v2 thruster operating at high 

anode efficiencies. As the figure shows, anode efficiencies achieved peak values between 50 and 70 A for 

all discharge voltages tested. At discharge currents below this peak value range, anode efficiencies 

decreased with decreasing discharge current because electron and neutral number densities within the 

discharge channel were likely too low to support efficient ionization of the propellant. This also occurred 

during operation at high magnetic field strengths, as shown in Figure 9 for the laboratory model thruster. 

However, at discharge currents above 50 to 70 A for operation of the v2 thruster at high anode 

efficiencies, anode efficiencies decreased with increasing current likely because charged particle and 

neutral number densities within the discharge channel were too high and this led to enhanced electron 

mobility to the anode. For the laboratory model thruster at high magnetic fields, anode efficiencies either 

continued to increase with increasing discharge current or achieved plateau values for discharge voltages 

 300 V. It is speculated that this was due to the higher magnetic fields, which led to a higher Hall 

parameter and therefore reduced electron mobility and diffusion to the anode. This reduction of electron 

current to the anode at high magnetic fields is likely the reason that the laboratory model thruster was able 

Figure 8.—Anode efficiency as a function of discharge current at various 

discharge voltages for the NASA-457M v2 during operation in high anode 

efficiency mode. 

0.55

0.57

0.59

0.61

0.63

0.65

0.67

0.69

0.71

0.73

10 30 50 70 90 110

Anode Efficiency

Discharge Current, A

500 V

400 V

300 V

250 V

200 V
NASA/TM—2012-217711 11 

Figure 9.—Anode efficiency as a function of discharge current at various 

discharge voltages for the laboratory model NASA-457M during operation 

with high magnetic fields. Data are from References 7 and 8. 

to sustain a lower discharge current for a given anode mass flow at discharge currents  60 A than that of 

the v2 thruster at lower magnetic fields, as shown in Figure 7. The behavior depicted in Figure 8 is not 

unique to the NASA-457M v2. It also occurred, for example, with the NASA-400M and BHT-20K at 

several discharge voltages (Refs. 17 and 18).

Cathode Performance 

Cathode keeper voltage as a function of discharge current is plotted in Figure 10 for both operating 

modes. As the figure demonstrates, cathode keeper voltages were not sensitive to discharge voltages or 

operating mode and were within 2.1 to 6.2 V throughout the performance test. This was not the case for 

cathode-to-ground voltages as a function of discharge current for both operating modes, as shown in 

Figure 11. During operation with high anode efficiency, cathode potentials were always within 13 V of 

facility ground. However, during operation at high magnetic fields, cathode-to-ground voltage magnitudes 

increased dramatically, exceeding 30 V, and showed a sensitivity to discharge voltage. 

The impacts of higher cathode-to-ground voltage magnitudes during high magnetic field operation 

has already been explored by Manzella et al., who investigated the impact of cathode mass flow on 

cathode-to-ground voltages and anode efficiencies for the laboratory model thruster (Ref. 7). They found 

that decreasing cathode mass flows and increasing discharge currents led to increasing cathode-to-ground 

voltage magnitudes. In addition to this, increased cathode-to-ground voltage magnitudes led to decreased 

anode efficiencies. They speculated that this was due to the development of large coupling voltage, which 

represented a thrust loss mechanism (Ref. 7). 

A comparison of the cathode-to-ground voltages measured with the v2 thruster at high magnetic 

fields to those of the laboratory thruster uncovered two findings: 

1. The increasing cathode-to-ground voltage with decreasing discharge current in Figure 11 was due to a 

deceasing cathode mass flow. 

2. Both thrusters exhibited similar cathode-to-ground voltage magnitudes along with similar sensitivities 

to cathode mass flow and discharge current. 

This confirmed that both thrusters operated similarly during operation with high magnetic fields. 

0.35

0.40

0.45

0.50

0.55

0.60

0.65

10 30 50 70 90 110

Anode Efficiency

Discharge Current, A

500 V

400 V

300 V

200 V
NASA/TM—2012-217711 12 

Figure 10.—Cathode keeper voltage as a function of discharge current for the 

NASA-457M v2 during operation in high anode efficiency and high magnetic 

field modes. 

Figure 11.—Cathode-to-ground voltage as a function of discharge current for 

the NASA-457M v2 during operation in high anode efficiency and high 

magnetic field modes. Cathode mass flows were 8 percent of the anode 

mass flows. 

 

0

1

2

3

4

5

6

7

10 30 50 70 90 110

Cathode Keeper Voltage,

V

DischargeCurrent, A

High Anode Efficiency

High Magnetic Field

‐35

‐30

‐25

‐20

‐15

‐10

‐5

0

10 30 50 70 90 110

Cathode‐to‐Ground Voltage,

V

DischargeCurrent, A

High Anode Efficiency

200 V ‐ High Magnetic Field

400 V ‐ High Magnetic Field
NASA/TM—2012-217711 13 

It was unclear, however, why the cathode-to-ground voltage sensitivity to cathode flow and discharge 

current changed so dramatically between the two operating modes. To uncover the root cause, the axial 

magnetic field strength in the immediate vicinity of the centrally-mounted cathode was measured as a 

function of the peak radial magnetic field along the discharge channel centerline. The axial magnetic 

fields in the vicinity of the cathode in the v2 and laboratory model thrusters were found to be directly 

proportional to the radial field in the discharge channel and nearly identical for the two thrusters. More 

importantly, this axial magnetic field strength at the cathode assembly exit plane was 65 percent of the 

peak radial magnetic field along the discharge channel centerline. When the v2 thruster was operated with 

high anode efficiencies, the axial magnetic field at the cathode assembly exit plane was significantly less 

than that during high magnetic field operation. In addition to this, the axial magnetic field increased with 

increasing axial distance from the cathode, achieving a peak value at about 4 cm downstream of the 

cathode assembly exit plane of both thrusters that was 20 to 30 percent larger than that at the cathode 

assembly exit plane. 

The impact of these strong axial magnetic fields in the vicinity of the cathode is speculated here. The 

cathode provides electrons to the discharge channel for ionization, ion acceleration, and neutralization of 

the outgoing ion beam. Positive potential gradients in the vicinity of the cathode assist with transporting 

electrons to the discharge and ion beam. Strong axial fields at the cathode can impede electron transport 

to either discharge or outgoing ion beam because electrons become trapped along magnetic field lines. 

Potential gradients (and therefore coupling voltages) must increase with increasing axial magnetic field 

strength to enable the required electron transport. At high cathode mass flows, however, increased 

electron-neutral collisions improve electron mobility to the discharge and ion beam, reducing coupling 

voltages, and therefore cathode-to-ground voltage magnitudes. 

The impact on coupling voltage is important because it is a thrust loss mechanism. And in the prior 

section, it was determined that the lower anode efficiencies of high magnetic field operation were due 

largely to lower measured thrusts. So quantifying its impact on thrust could give insight to the cause of 

the reduced thrust at high magnetic fields. Unfortunately, this effect was not studied in detail during this 

performance test. However, for one operating point at high magnetic field strength (i.e., for a discharge 

current and voltage of 50 A and 400 V, respectively), increasing the cathode flow decreased the cathodeto-ground voltage magnitude by 10 V to a level that was comparable to that of high anode efficiency 

operation. This cathode-to-ground voltage magnitude decrease increased the thrust by about 2 percent. 

However, this thrust improvement was still 10 percent lower than that measured during high anode 

efficiency operation. This datum suggests that while coupling voltage did play a role in reducing thrust 

during high magnetic field operation, it did not account for all of the disparity. To fully account for the 

thrust disparity, the contributions of multiply charged ions and ion beam divergence to thrust losses 

would have to be measured. 

Overall Thruster Performance 

Thrust and thrust-to-power as a function of total thruster input power are plotted in Figures 12 and 13, 

respectively, for the v2 thruster operating with high anode efficiency. The uncertainty in total power and 

thrust-to-power for all results reported herein were estimated to be 0.3 and 2.0 percent, respectively, of 

the indicated values. The peak thrust demonstrated by the NASA-457M v2 was 2.3 N and this occurred at a 

total thruster input power of 50.0 kW. The thrust was nonlinear as a function of total input power at the 

higher power levels at each discharge voltage, which was unlike the laboratory model thruster operating 

with high magnetic fields (Refs. 7 and 8). This was because the total input power was dominated by 

discharge power. The cause for the nonlinear behavior of thrust as a function of discharge current for a 

given discharge voltage has already been discussed (see the discussions for Figs. 6 and 7). As shown in 

Figure 13, the thruster demonstrated a thrust-to-power range of 76.4 mN/kW at low power to 46.1 mN/kW 

at full power. 
NASA/TM—2012-217711 14 

Figure 12.—Thrust as a function of total thruster input power at various discharge 

voltages for the NASA-457M v2 thruster operating with high anode efficiency. 

Figure 13.—Thrust-to-power as a function of total thruster input power at various 

discharge voltages for the NASA-457M v2 thruster operating with high anode 

efficiency. 

Specific impulse as a function of total thruster input power is shown in Figure 14 for the NASA457M v2 thruster operating with high anode efficiency. The uncertainty in specific impulse was estimated 

to typically be 2.3 percent of the indicated value. Direct comparisons with the laboratory model thruster 

were not possible because the laboratory model thruster was operated at fixed cathode flows rather than 

the 8 percent of anode mass flow used by the v2 thruster. As shown in Figure 14, the v2 thruster 

demonstrated specific impulse range of 1420 to 2740 s. The figure also shows that the v2 thruster was 

able to operate at a specific impulse of about 2000 s with a discharge voltage of only 300 V. This is 

significant because the HEFT trade studies baselined a thruster specific impulse of about 2000 s and 

lower discharge voltages are typically thought to lead to longer thruster service life capability (Ref. 2). 

0

500

1000

1500

2000

2500

0 10,000 20,000 30,000 40,000 50,000

Thrust, mN

Total Input Power, W

200 V

250 V

300 V

400 V

500 V

45

50

55

60

65

70

75

80

0 10,000 20,000 30,000 40,000 50,000

Thrust‐to‐Power, mN/kW

Total Input Power, W

200 V

250 V

300 V

400 V

500 V
NASA/TM—2012-217711 15 

Figure 14.—Specific impulse as a function of total thruster input power at 

various discharge voltages for the NASA-457M v2 thruster operating with 

high anode efficiency. 

Figure 15.—Thrust efficiency as a function of total thruster input power at 

various discharge voltages for the NASA-457M v2 thruster operating with 

high anode efficiency. 

The thrust efficiency as a function of total thruster input power is shown in Figure 15 for the 

NASA-457M v2 thruster operating with high anode efficiency. The uncertainty in thrust efficiency was 

estimated to typically be 4.2 percent of the indicated value. As before, direct comparisons with the 

laboratory model thruster were not possible because it was operated at fixed cathode flows rather than the 

8 percent of anode mass flow used by the v2 thruster. Regardless, thrust efficiency is dominated by anode 

performance, and anode efficiencies comparisons were made between the two thrusters in an earlier 

section. Thrust efficiencies for the v2 thruster varied from a low of 0.51 at low power to 0.66 at 30 kW 

with a discharge voltage of 500 V. For a discharge voltage of 300 V, where specific impulses would be 

about 2000 s, thrust efficiencies varied from 0.57 to 0.63. 

1200

1400

1600

1800

2000

2200

2400

2600

2800

0 10,000 20,000 30,000 40,000 50,000

Specific Impulse, s

Total Input Power, W

500 V

400 V

300 V

250 V

200 V

0.45

0.50

0.55

0.60

0.65

0.70

0 10,000 20,000 30,000 40,000 50,000

Thrust Efficiency

Total Input Power, W

500 V

400 V

300 V

250 V

200 V
NASA/TM—2012-217711 16 

Conclusions 

An excellent thruster candidate for a solar electric propulsion system that supports human exploration 

missions is the 50 kW-class NASA-457M Hall thruster. A laboratory model version had been 

successfully designed, built, and extensively tested, and the successful test campaign led to the 

development of a higher fidelity version of this thruster, labeled the NASA-457M v2. Testing of the 

NASA-457M v2 was re-initiated last fiscal year. Performance test results were reviewed in this report. 

The development goal for the NASA-457M v2 was to advance its technology readiness level to 4/5. 

The higher fidelity thruster retained the same critical discharge chamber dimensions, however the 

magnetic circuit was designed for improved performance and field symmetry. A laboratory model hollow 

cathode assembly was used for all performance testing reported herein. 

The thruster was mounted in the main volume of NASA GRC’s Vacuum Facility 5. Tank pressures 

never exceeded 4.0×10–3 Pa (3.0×10–5 torr) during performance testing. Xenon was fed to the thruster 

through calibrated mass flow controllers and power was provided by commercially available power 

supplies. A null-type inverted pendulum thrust stand was used to measure thrusts. 

Thruster discharge power was varied from 5 to 50 kW over discharge voltage and current ranges of 

200 to 500 V and 15 to 100 A, respectively. Anode efficiencies during high efficiency operation for the 

NASA-457M v2 varied from 0.56 to 0.71. The peak efficiency was similar to that of other state-of-the-art 

high power Hall thrusters, but outperformed these other thrusters at lower discharge voltages. A 

comparison of the v2 and laboratory model thruster anode efficiencies showed two differences: 1) the v2 

thruster significantly outperformed the laboratory model thruster; and 2) the v2 thruster exhibited a 

decreasing efficiency with increasing power at high discharge powers. 

The main reason for the anode efficiency differences between the two thruster versions was 

differences in how they were operated. The laboratory model thruster was operated with a nearly constant 

peak centerline radial magnetic field. During v2 thruster testing, two stable discharge modes of operation 

were found. One was at low magnetic field strengths, which produced high anode efficiencies, and the 

other at high magnetic fields. When the v2 thruster was operated at a constant peak radial magnetic field 

that was 75 percent that of the laboratory model thruster, anode efficiencies decreased between 0.05 and 

0.11, with the performance of the v2 thruster approaching that of the laboratory model. This indicated that 

anode performance between the two thruster versions was actually quite similar and that the NASA-457M 

v2 could be operated at lower magnetic fields to achieve more optimal anode efficiencies than had been 

done in the past. The reason for the lower anode efficiencies at high magnetic fields was that operation in 

this mode produced lower thrusts, even when the data suggested that higher magnetic fields were 

producing higher beam ion currents. This suggests that other thrust loss mechanisms are becoming 

significant during operation with high magnetic fields, such as beam divergence, the ratio of multiply 

charged ions, or coupling voltage losses. 

The anode efficiency trends with discharge power for high anode efficiency operation were a function 

of discharge current. Anode efficiencies achieved peak values between 50 and 70 A for all discharge 

voltages tested. At discharge currents above peak values, anode efficiencies decreased with increasing 

current because charged particle and neutral number densities within the discharge channel were too high 

and this led to enhanced electron mobility to the anode. The laboratory model thruster did not exhibit this 

trend likely because it was operated at a high magnetic field strength, and therefore a reduced electron 

mobility and diffusion to the anode. 

Cathode keeper voltages were not sensitive to discharge voltages or mode of operation and were 

within 2.1 to 6.2 V throughout the performance test. During operation with high anode efficiency, cathode 

potentials were always within 13 V of facility ground. However, during operation at high magnetic fields, 

cathode-to-ground voltage magnitudes increased dramatically, exceeding 30 V, which was similar to the 

laboratory model thruster behavior. The root cause was the axial magnetic field strength in the immediate 

vicinity of the centrally-mounted cathode, which were found to be directly proportional to the radial field 

in the discharge channel and nearly identical for the two NASA-457M thrusters. More importantly, this 

axial magnetic field strength was 65 percent of the peak radial magnetic field along the discharge channel 
NASA/TM—2012-217711 17 

centerline. The speculated impact of these strong axial magnetic fields in the vicinity of the cathode was 

to increase coupling voltages and, therefore, decrease thrust. Test data suggested that this thrust loss was 

indeed occurring, however, it did not fully account for the lower thrust during operation with high 

magnetic fields. 

The peak thrust demonstrated by the NASA-457M v2 was 2.3 N at a total thruster input power of 

50.0 kW. The thruster also demonstrated a thrust-to-power range of 76.4 mN/kW at low power to 

46.1 mN/kW at full power and a specific impulse range of 1420 s to 2740 s. It was able to operate at a 

specific impulse of about 2000 s with a discharge voltage of only 300 V. This is significant because the 

HEFT trade studies baselined a thruster specific impulse of about 2000 s and lower discharge voltages are 

typically thought to lead to longer thruster service life capability. For a discharge voltage of 300 V, thrust 

efficiencies varied from 0.57 to 0.63. 

These performance test results have shown that the NASA-457M v2 thruster is a potential candidate 

for future high power manned missions. This thruster design, or a derivative of it, could provide NASA 

with a potential head start for future human space exploration. 

References 

1. “NASA Exploration Systems Mission Directorate presentation, Human Space Exploration 

Summary,” http://www.nasa.gov/exploration/new_space_enterprise/home/heft_summary.html, 

January 2011. 

2. Mercer, C.R., et al., “Benefits of Power and Propulsion Technology for a Piloted Electric Vehicle to 

an Asteroid,” AIAA Paper 2011-7252, September 2011. 

3. Capadona, L.A., et al., “Feasibility of Large High-Powered Solar Electric Propulsion Vehicles: 

Issues and Solutions,” AIAA Paper 2011-7251, September 2011. 

4. Brophy, J.R., et al., “300-kW Solar Electric Propulsion System Configuration for Human 

Exploration of Near-Earth Asteroids,” AIAA Paper 2011-5514, August 2011. 

5. Hofer, R.R. and Randolph, T.M., “Mass and Cost Model for Selecting Thruster Size in Electric 

Propulsion Systems,” AIAA Paper 2011-5518, August 2011. 

6. Dunning, J. and Sankovic, J., “NASA’s Electric Propulsion Program,” AIAA Paper 2000-3145, 

July 2000. 

7. Manzella, D.M., Jankovsky, R., and Hofer, R.R., “Laboratory Model 50 kW Hall Thruster,” 

AIAA Paper 2002-3676, July 2002. 

8. Manzella, D.M., “Scaling Hall Thrusters to High Power,” Ph.D. Dissertation, Department of 

Mechanical Engineering, Stanford University, Stanford, CA, June 2005. 

9. Jankovsky, R.S., et al., “NASA’s Hall Thruster Program 2002,” AIAA Paper 2002-3675, July 2002. 

10. Jacobson, D.T. and Manzella, D.M., “50 kW Class Krypton Hall Thruster Performance,” 

AIAA Paper 2003-4550, July 2003. 

11. Manzella, D. and Jacobson, D., “Investigation of Low-Voltage/High-Thruster HALL Thruster 

Operation,” AIAA Paper 2003-5004, July 2003. 

12. Jacobson, D.T., et al., “NASA’s 2004 Hall Thruster Program,” AIAA Paper 2004-3600, July 2004. 

13. Kamhawi, H., et al., “Overview of Hall Thruster Activities at NASA Glenn Research Center,” 

32nd International Electric Propulsion Conference, IEPC Paper 2011-339, September 2011. 

14. Hofer, R.R., Peterson, P.Y., and Gallimore, A.D., “A High Specific Impulse Two-Stage Hall 

Thruster with Plasma Lens Focusing,” IEPC Paper 01-036, 27th International Electric Propulsion 

Conference, October 2001. 

15. Hofer, R.R., Gallimore, A.D., “The Role of Magnetic Field Topography in Improving the 

Performance of High-Voltage Hall Thrusters,” AIAA-2002-4111, July 2002. 

16. Kamhawi, H., et al., “Performance Evaluation of the NASA-300M 20 kW Hall Effect Thruster,” 

AIAA Paper 2011-5521, August 2011. 
NASA/TM—2012-217711 18 

17. Peterson, P.Y., Jacobson, D.T., Manzella, D.H., and John, J.W. “The Performance and Wear 

Characterization of a High-Power High-Isp NASA Hall Thruster,” AIAA Paper 2005-4243, 

July 2005. 

18. Szabo, J., et al., “A Commercial One Newton Hall Effect Thruster for High Power In-Space 

Missions,” AIAA Paper 2011-6152, August 2011. 

19. Kim, V., “Main Physical Features and Processes Determining the Performance of Stationary Plasma 

Thrusters,” Journal of Propulsion and Power, vol. 14, no. 5, September-October 1998, pp 736-743. 

20. Huang, W., et al., “Ion Current Density Study of the NASA-300M and NASA-457M v2 Hall 

Thrusters,” to be presented at the 48th Joint Propulsion Conference and Exhibit, July-August 2012. 

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Performance Test Results of the NASA-457M v2 Hall Thruster 

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Soulas, George, C.; Haag, Thomas, W.; Herman, Daniel, A.; Huang, Wensheng; Kamhawi, 

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14. ABSTRACT

Performance testing of a second generation, 50 kW-class Hall thruster labeled NASA-457M v2 was conducted at the NASA Glenn Research 

Center. This NASA-designed thruster is an excellent candidate for a solar electric propulsion system that supports human exploration 

missions. Thruster discharge power was varied from 5 to 50 kW over discharge voltage and current ranges of 200 to 500 V and 15 to 100 A, 

respectively. Anode efficiencies varied from 0.56 to 0.71. The peak efficiency was similar to that of other state-of-the-art high power Hall 

thrusters, but outperformed these thrusters at lower discharge voltages. The 0.05 to 0.18 higher anode efficiencies of this thruster compared 

to its predecessor were primarily due to which of two stable discharge modes the thruster was operated. One stable mode was at low 

magnetic field strengths, which produced high anode efficiencies, and the other at high magnetic fields where its predecessor was operated. 

Cathode keeper voltages were always within 2.1 to 6.2 V and cathode voltages were within 13 V of tank ground during high anode 

efficiency operation. However, during operation at high magnetic fields, cathode-to-ground voltage magnitudes increased dramatically, 

exceeding 30 V, due to the high axial magnetic field strengths in the immediate vicinity of the centrally-mounted cathode. The peak thrust 

was 2.3 N and this occurred at a total thruster input power of 50.0 kW at a 500 V discharge voltage. The thruster demonstrated a thrust-topower range of 76.4 mN/kW at low power to 46.1 mN/kW at full power, and a specific impulse range of 1420 to 2740 s. For a discharge 

voltage of 300 V, where specific impulses would be about 2000 s, thrust efficiencies varied from 0.57 to 0.63. 

15. SUBJECT TERMS

Hall thruster; Stationary plasma thruster 

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